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THE AIRPLANE ENGINE 




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Mn»r^^ 




THE 

AIRPLANE ENGINE 



BY 



LIONEL S. MARKS, B.Sc, M.M.E. 

PROFESSOR OP MECHANICAL ENGINEERING, HARVARD UNIVERSITY 

MEMBER AMERICAN SOCIETY MECHANICAL ENGINEERS, 

FELLOW AMERICAN ACADEMY ARTS AND SCIENCES 



First Edition 



McGRAW-HILL BOOK COMPANY, Inc. 
NEW YORK: 370 SEVENTH AVENUE 

LONDON : 6 & 8 BOUVERIE ST., E. C. 4 

1922 






McGraw-Hill Book Company, Inc. 



JAN 2 1 (922 



THE MAPI.E I'KKSS YORK PA 



£?CU653618 



w , ^W 






tN 



ZJo 

MY WIFE 

Josephine Preston Peabody 



PREFACE 

This volume attempts two things: to formulate existing 
knowledge of the functioning of the airplane engine and its 
auxiliaries; and to present and discuss the essential constructive 
details of those engines whose excellence has resulted in their 
survival. 

The material here collected is largely new; very little of it 
could have been written before the war and only a small frac- 
tion was available for publication before 1919. It is based 
mainly on the researches and engine developments originating 
during the war and resulting from the war's urgencies. The 
researches have been carried out almost exclusively under 
governmental auspices; in the United States at the Bureau of 
Standards and at the Air Service experimental plant at McCook 
Field; in Great Britain at the Royal Aircraft Factory and the 
National Physical Laboratory; in France and Germany at 
equivalent institutions. Many of the results of these investiga- 
tions were published confidentially during the war in Reports of 
the Bureau of Standards; in Bulletins and Technical Orders of 
the Airplane Engineering Division of the U. S. Army; in Reports 
of the (British) Advisory Committee for Aeronautics; in Bulle- 
tins de la Section Technique de l'Aeronautique Militaire; and 
in Technische Berichte. This material has now become avail- 
able and much of it has been published in the Reports of the 
(U. S.) National Advisory Committee for Aeronautics and in 
the technical press. 

Similarly, the constructive details of most of the existing 
airplane engines are now available, chiefly from descriptions of 
captured machines. The German and Austrian engines captured 
by the British were subjected to a technical analysis which has set 
a new standard in such matters. Not only were the engines and 
their auxiliaries tested exhaustively for performance but all the 
parts were minutely measured, loads and stresses calculated, 
and the metal analyzed for chemical composition. The French 
carried out similar analyses of German engines. The Germans 
published corresponding, though less detailed, analyses of 

vii 



viii PREFACE 

French, English and American engines. Since the war, the 
U. S. Air Service has also analyzed American and foreign engines 
and has published its findings in Technical Orders and Informa- 
tion Circulars. 

With all this material, the designer of the airplane engine 
has at hand more detailed precedent from which to depart than 
is available for other types of engine. 

The writer desires to acknowledge his indebtedness to Professor 
E. B. Warner and to Lieut. E. E. Aldrin for assistance in obtaining 
information; and to Mr. R. H. Taylor for assistance in reading 
proofs and in preparing the index. 

L. S. M. 

Cambridge, Mass. 
January, 1922. 



CONTENTS 



Page 

Prefaue vii 

Chapter 

I. Power Required and Power Available 1 

II. Engine Efficiencies and Capacities 11 

"^ III. Engine Dynamics .40 

IV. Engine Dimensions and Arrangements 60 

V. Materials 114 

VI. Engine Details 122 

VII. Valves and Valve Gears . 151 

VIII. Radial and Rotary Engines 176 

~ IX. Fuels and Explosive Mixtures 212 

X. The Carburetor 245 

XI. Fuel Systems 289 

XII. Ignition 295 

XIII. Lubrication 327 

^XIV. The Cooling System 344 

. XV. Geared Propeller Drives 378 

XVI. Supercharging 386 

XVII. Manifolds and Mufflers 416 

XVIII. Starting 424 

XIX. Potential Developments 434 

Index 445 



IX 



THE AIRPLANE ENGINE 

CHAPTER I 

POWER REQUIRED AND POWER AVAILABLE 

Power Required for Flight. — An airplane in flight is sustained 
by the lift of the wings. Consider the wing as a thin flat plate, 
Fig. .1. Four forces are acting: 




Fig. 1. — Forces acting on a flat plate. 

1. The weight, W, of wing and parts supported thereby, 
downward. 

2. The thrust, T, or forward impulse due to the propeller. 

3. The lift, L, of the air which acts in a direction perpendicular 

1 



2 THE AIRPLANE ENGINE 

to that of the plane with respect to the air, and produces 
sustentation. 

4. The wing resistance or drag, D, measured perpendicularly 
to L, the component of the total force on the wing which opposes 
forward motion. 

At constant horizontal speed, L = W and D = T. As these 
pairs of forces are not acting in the same lines they give rise to 
turning moments and it is necessary for stability that L X m = 
T X n. Both L and D are created by the velocity of the plane; 
the direction of the latter is concurrent with (but opposite to) 
the path of flight. The former is directed at right angles with 
the path of flight. In horizontal flight L and D are vertical and 
horizontal forces respectively. 

Wing Characteristics. — Flat Plates. — A plate moving with 
respect to the air undergoes an approximately normal pressure 
F which is proportional to the density of the air and (within 
limits) to the square of the relative velocity. This pressure may 
be resolved into components L and D perpendicular and parallel 
to the path of flight. F 2 = L 2 + D 2 . The component L is 
useful, while D is objectionable. The pressure F depends on the 
area of the plate, but varies somewhat with its shape. 

The incidence, i, or angle between the plate and the flight 
path has a dominating influence on the resulting pressure; and 
particularly on the relation between L and D. Since L = F cos i 
and D = F sin i, L/F decreases and D/F increases, as i is 
increased from deg. 

For a given angle of incidence the following relations hold: 

L = K L dAV 2 
D = K D dAV 2 

where K L and K D are experimentally determined lift and drag 
coefficients respectively; d is the air density; A is the wing area; 
and V the relative air velocity. The usual units are L and D in 
pounds; d, relative air density in terms of normal density; A in 
square feet; and V in miles per hour. From the above equations 

it is seen that jr= jsr* That is, lift is obtained with minimum 
drag whenv^- is a maximum. 

It is found that more favorable ratios of K L /K D are obtained 
with curved wings, as in Fig. 2, than with flat plates. The angle 



POWER REQUIRED AND POWER AVAILABLE 3 

of incidence of such a wing is arbitrarily denned as the acute 
angle between the wind direction and the lower chord of the wing. 
Figure 3 gives the values of K L , K D , and K L /K D or L/D for the 
wing or aerofoil section shown in Fig. 2, and shows a maximum 
value of L/D of 17 at an angle of incidence of 3 deg. 



ChoroP r 
Fig. 2. — Cross section of wing. 

An airplane consists not only of the wings which give susten- 
tation, but also of other members such as the fuselage, radiator, 
landing gear, and wing bracing. These give no aid in sustaining 
the plane but offer a resistance, the parasite resistance, which 



0.0030 



0.0024 



0.0020 


20 


0.0016 


16 


:* 0.00 12 


12 


jjo 


0.0008 


8 


0.0004 


4 







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*K L 






\ 














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/ 






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1/ 






















■K D 




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1 
















1 

















0.0008 



0.0006 



-8-4 4 8 12 16 20 

Angle of Incidence , Degrees 

Fig. 3. — Lift and drag coefficients of a wing. 



0.0004 



must be overcome in flight. If P is the parasite resistance in 
pounds 

P = K N dV 2 

where K N is a coefficient with practically constant value in any 
given airplane. 



4 THE AIRPLANE ENGINE 

The total resistance, R, to the motion of the plane in the direc- 
tion of flight is the sum of the wing and parasite resistances, or 

R = D+P = d{K D A + K N )V 2 

This resistance must equal the propeller thrust, T, in uniform 
flight. The horse power required to overcome the resistance is 
_ 1.467 fl7 = RV 
R " 550 375 

where 1.467 is the constant to convert miles per hour to feet 
per second and 550 is the equivalent of 1 hp. in foot-pounds per 
second. If the efficiency of the propeller is e, the brake horse 
power, H B , required at the engine for horizontal flight at speed V 
is: 

RV_ = d(K D A + K N )V* 
B 375e 375e 

The lift, L, must equal the weight of the plane, W, in horizontal 
flight, and the equation W = K L dAV 2 must be satisfied simul- 
taneously with the b.h.p. equation, with values of K L and K D 
corresponding to some one angle of incidence. In a design 
W, A, V, K N are known by assumption or calculation and it is 
required to find H R , K L and K D for a series of values of V. Since 
K L and K D are functions of i there are only two independent 
variables H R and i to be found. 

The power required in an airplane of 200 sq. ft. wing area 
with a total weight of 1,200 lb. using the wing with the properties 
shown in Fig. 3 and with parasite resistance P = 0.0025 V 2 
is shown in Fig. 4. It will be seen that the plane has a minimum 
possible velocity of about 44 miles per hour but that the power 
required to drive it starts to increase very rapidly if the speed 
gets below 46 miles per hour. The flying conditions are for 
speeds of 46 miles per hour or higher. At lower speeds there are 
reversed controls, that is, more power is required to go slower, 
and the angle of incidence is so high that the plane will be in 
danger of stalling. The power required curve of Fig. 4 is of a form 
which may be regarded as typical. 

If the propeller does more work than is required the plane will 
climb and its rate of climbing will depend only on the amount of 
the excess power; all the additional work goes into raising the 
plane. If, on the other hand, power is deficient and the propel- 
ler does less work than is required to keep the plane in level 
flight at the existing speed, the plane will glide down and the work 



POWER REQUIRED AND POWER AVAILABLE 



done by the falling plane (weight X fall) will be exactly equal 
to the difference between the work required to keep the plane 
moving with its existing speed in the direction of flight and the 
work done by the propeller. If the power available is at any 
time in excess of that required for level flight, it can be reduced by 
throttling the engine. There is only one speed possible in 
level flight for any given angle of incidence. Opening the throttle 
does not, as with automobiles, increase the speed in level flight, 
but will start the airplane climbing unless the incidence is also 

120 



100 



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40 50 60 70 80 90 

Velocity, Miles per Hour 

Fig. 4. — Characteristics of an airplane in level flight 



100 



changed by operating the elevator. The power supplied by the 
engine with a given incidence cannot affect the velocity of the plane 
but will determine whether the plane climbs, flies level, or glides. 

TV 

The horse power available for driving the plane is given by w=z, 

that is, at any plane speed it depends only on the propeller 
thrust. As the thrust depends on the engine speed both the 
engine and the propeller characteristics must be examined before 
the thrust can be determined. 

Air Propellers. — The propeller converts the torque at the 
engine crankshaft into a thrust along the axis 'of the propeller 
shaft. Its action is remotely similar, to the turning of a solid 



6 THE AIRPLANE ENGINE 

screw in a solid nut and the same general terminology is em- 
ployed. The axial distance that would be travelled by the pro- 
peller for one revolution if the air were incompressible is called 
its nominal pitch; the nominal pitch divided by the outside 
diameter is the nominal pitch ratio. An air propeller does not, 
however, advance a constant amount for one revolution; it is 
advancing in a medium which is readily displaced, and it is 
found that it must advance each revolution a distance greater 
than the nominal pitch if it is not to displace the air backward. 
This greater distance is called the dynamic pitch. The difference 
between the dynamic pitch and the actual or effective pitch is called 
the slip. Positive thrust can be obtained only when there is slip 
and the greater the slip the greater will be the thrust; maximum 
thrust is obtained when the plane is on the ground in which case 
the effective pitch is zero. 

If a propeller makes N revolutions per minute and the plane 
moves through the air with a velocity of V feet per minute the 
effective pitch is V/N and the effective pitch ratio is V/ND. 

The characteristics of propellers are most easily determined by 
tests on models. It is found that geometrically similar propellers 
have the same characteristics for the same values of V/ND. 
The important characteristics are the thrust, T, lb.; torque, 
Q, lb.-ft.; torque horse power, H B ; thrust horse power, H R ; and 
efficiency, e. The torque horse power is necessarily the same as 
the engine brake horse power. The efficiency, e = H R /H B > 

Torque and thrust are given by the following equations: 

w V 2 D* • a 



T = 



1000 
w V 2 D 2 



100 

where a and b are the torque and thrust coefficients respectively 
and w is the air density in pounds per cubic foot. 

Efficiency is obtainable directly from these coefficients. 

Efficiency = 

Thrust horse power _ TV = 10b V = b V 
Torque horse power " 2rQN ~~ 2wa ND " a ND 

Values of these coefficients and of e have been determined by 
Durand 1 for propellers of many types and proportions. The 

1 Nat. Adv. Comm. on Aeronautics, 1917. 



POWER REQUIRED AND POWER AVAILABLE 



curves of Fig. 5 give typical values obtained from three propel- 
lers which differ only in nominal pitch ratio; the values of the 
pitch ratio are 0.5, 0.7 and 0.9 respectively. By the use of such 
curves the horse power required can be easily obtained. For 
example, assume, a propeller 8 ft. in diameter (Fig. 5) with pitch 
ratio 0.9, making 1,200 r.p.m. and with a speed of 72 miles per 
hour (105.6 ft. per second) at an elevation of 3,000 ft. The value 




Fig. 5. — Propeller coefficients and efficiencies. 

of V/ND = 0.66. From Fig. 5,6 = 0.685, and a = 0.970. Also 
w = 0.071 lb. per cubic foot. Then 

0.071 X (105.6) 2 X 8 2 X 0.685 



T = 



100 



= 347 lb. 



n - 0-071 X (105.6) 2 X 8 3 X 0.970 _ QQ „ 



Torque horse power, H B , = 

2wQN = 2tt X 393 X 1,200 
33,000 



= 89.5 



33,000 
Efficiency = 1.59 X jj^ X 0.66 = 0.742 



8 THE AIRPLANE ENGINE 

The efficiency can also be read directly from Fig. 5. 

Power Available for Flight. — The horse power available with a 
given engine and given plane speed, V, can now be determined. 
Throughout a considerable range of revolutions per minute the 
mean effective pressure in the engine is practically constant but 
falls off at highest speeds. If it is assumed constant the propeller 
torque will be constant. The torque equation can be written 

a ( V y = Q 
1,000 ' \ND/ ' N 2 D 5 

or since Q and D are constant Jj 

Constant 



N = 



(Id) x Va 



v 

Assuming various values of v^ and finding in Fig. 5 the 

corresponding values of a, a series of values of N can be obtained 
and substituting these values in V/ND the corresponding values 
of V are determined. That is, the change of revolutions per 
minute with flying speed, V, or with V/ND is obtained. In 
Fig. 6 there are plotted curves 1 showing this variation for 
propellers of nominal pitch ratios of 0.5, 0.7 and 0.9 respectively. 
These curves do not correspond exactly to the coefficients of Fig. 
5; they average the results of Durand's tests and apply fairly 
well to standard forms of propeller. The unit values of V/ND 
and N are those corresponding to maximum efficiency. The 
product of the ordinates and abscissae for any point is (V/ND) 
X N = V/D, and since D is constant, the ratio of this product 
at'' two points is also the ratio of the corresponding plane 
velocities. 

As an example suppose a propeller with nominal pitch ratio 
0.7 designed for an airplane flying normally at 100 miles per hour 
and that the full engine power is absorbed at 1,600 r.p.m. If 
the speed of the machine increases so that V/ND = 1.1, then 
N becomes 1.052 and V increases in the ratio 1.1 X 1.052 = 
1.157. The propeller will then turn 1.052 X 1,600 = 1,683 r.p.m. 
on full throttle at a plane speed of 115.7 miles per hour. 

The torque horse power is proportional to the engine speed; 
the efficiency can be obtained from Fig. 5, and multiplied by 

1 Supplied by E. P. Warner. 



POWER REQUIRED AND POWER AVAILABLE 



9 



the torque horse power will give the available horse power. In 
the example just given, assume that the propeller has the char- 
acteristics shown in Fig. 5 and that the torque horse power at 
maximum efficiency is 100. Maximum efficiency is seen to be 
for V/ND = 0.64, which therefore corresponds to the unit of 
Fig. 6. For V/ND = 1.1 on Fig. 6 we have V/ND = 0.64 X 
1.1 = 0.704 in Fig. 5; the corresponding efficiency is seen to be 
0.74. The torque horse power is 100 X 1.052 and the available 
horse power is 105.2 X 0.74 = 78. By finding the available 























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1 
1 

t 






















V 

— n 4- — 


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0.4 



0.6 



0.8 



V/ND 



1.4 



Fig. 6. — Variation of revolutions per minute N, with speed of airplane, V; 
engine torque constant. 



horse power for a number of other values of V/ND and plotting 
them against V the available horse power curve of Fig. 4 is 
obtained. 

If engine torque varies with the engine speed, N, another pro- 
cedure must be followed in finding the values of N corresponding 
to different values of V. Taking the propeller torque equation 

Q = a series of curves may be drawn, as in Fig. 7, each 

giving the change in propeller torque with speed at constant V. If 
the engine torque is now drawn on the same figure its intersection 



10 



THE AIRPLANE ENGINE 



with the constant V curves will give the values of N at which 
engine and propeller torques are equal, that is, it will give the 
operating speeds. The available horse power can then be 
found by the use of the efficiency curve of Fig. 5 

Returning to Fig. 4 it is seen that with wide-open throttle the 
available horse power is equal to the required horse power at 45 
miles per hour and at 90 miles per hour; between these limits the 
available horse power is in excess of the power required for level 



100 



600 



500 



400 



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200 



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1100 



1200 



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1400 1500 

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1600 



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1600 



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Fig. 7. — Variation of engine and propeller torque with revolutions per minute 
at various airplane speeds. 

flight; outside these limits the engine pbwer is insufficient and the 
plane will glide. Between 45 and 90 miles per hour level flight 
can be maintained only by closing the throttle and the speed of 
flight will depend on the angle of incidence. With the throttle 
wide open the plane will climb and its rate of climb will be great- 
est at that speed and corresponding angle of incidence at which 
the difference between the available power and required power 
is a maximum. In the case of Fig. 4 this will be at about 65 
miles per hour and an angle of incidence of 3.6 deg. 



CHAPTER II 



ENGINE EFFICIENCIES AND CAPACITIES 



A typical airplane engine is shown in transverse section in 
Fig. 8 and in longitudinal section in Fig. 9. It has six vertical, 
single-acting, water-cooled cylinders in a row, driving the crank- 
shaft, b, through pistons, p, 
piston pins, g, connecting 
rods, r, and crankpins, k. 
The crankshaft is supported 
on seven bearings, carries 
the propeller hub, d, at its 
forward end with a thrust 
bearing, e, behind it and has 
a bevel wheel, /, at its rear 
end meshing with another 
bevel wheel on the vertical 
shaft, I. The shaft I drives 
the camshaft, c, through 
bevel gearing at its upper 
end and carries at its lower 
end the centrifugal water 
pump, n, and the two gear 
oil pumps, o; it also drives 
the magneto, m, through 
helical gears, q. The cam- 
shaft, c, carries inlet cams 
which act on the rocker 
levers, t, and open the inlet 
valves, v, against the com- 
pression of the valve springs, 
h; there is one inlet valve to 
each cylinder. The cam- 
shaft also carries cams which 
act directly on the tappets of the exhaust valves, u, of which 
there are two per cylinder. The water jackets, w, surround the 
cylinders and the valve cages. Air enters the carburetor, a, and 
goes through the inlet manifold, i, and past the inlet valve, v, to 

11 




Fig 



-Transverse section of Siddeley 
'Puma" airplane engine. 



12 



THE^AIRPLANE ENGINE 




ENGINE EFFICIENCIES AND CAPACITIES 



13 



JT 



[400 



\zoo 



.zzoo 



100 





d 


















































c 








a 


V 






== ="- e 

b 



Volumes ° 

Fig. 10. — Ideal indicator card 
of Otto-cycle engine. 



the cylinder, x. After compression the charge is ignited by the 
spark plugs, s, which get their current from the magneto, m. 

Actions Occurring in the Cylinder. — Figure 10 is a theoretic a 
indicator card for an engine using the Otto cycle, which is always 
used in aviation engines. The indicator card shows (vertically) 
the gas pressure inside the cylinder at each position of the piston. 
At the position a the piston is at the end of its stroke most remote 
from the crankshaft and the volume of is the volume of the clear- 
ance or combustion space; the total volume displaced by the 
piston is represented by fg and is the 
product of the piston area by its stroke; 
the maximum volume of gas in the cyl- 
inder is og. 

The cycle of operations inside the 
cylinder begins with the piston at a 
and with the gas inlet valve open, 
establishing free communication be- 
tween the cylinder and the external 
air. The piston makes its stroke from 
a to b with the inlet valve kept open; 

the pressure will remain atmospheric. This is the suction stroke. 
The inlet valve now closes and the piston returns. Since both 
valves are closed, the mixture in the cylinder is compressed (curve 
6c, compression stroke) the final pressure depending mainly upon 
the volume of the clearance space into which the gas is crowded at 
the inner end of the stroke. 

When the piston is at c (actually a little sooner than this) 
the compressed mixture is ignited and the pressure suddenly 
increases (line cd). This pressure drives the piston back, the 
burnt mixture expanding (curve de) while the valves remain 
closed. This is the power or expansion stroke. 

At the end, e, of the expansion stroke, the exhaust valve opens. 
The pressure drops to atmospheric (line eb). The piston returns 
and the burned gas is driven out at the exhaust valve at atmos- 
pheric pressure. At the end of this exhaust stroke, ba, the exhaust 
valve closes, the inlet valve opens, and the cycle recommences. 

SlJMMAKY 
Stroke Action Inlet valve Exhaust valve 

First out Suction Open Closed 

First in Compression Closed Closed 

Second out - Explosion and expansion Closed Closed 

Second in Exhaust Closed Open 



14 THE AIRPLANE ENGINE 

Pressures and Temperatures in the Ideal Cycle. — Let p be 

the absolute gas pressure in pounds per square inch, v the volume 
in cubic feet, t the temperature Fahrenheit and T the absolute 
temperature. The suffixes a, b, c, etc., indicate the points on the 
cycle, shown in Fig. 10, at which these quantities are being 
considered. In the ideal cycle, the volume v b — v a of explosive 
mixture is taken into the cylinder during the admission period 
and is still at atmospheric pressure and temperature at the end 
of the stroke b. The mixture or " charge" is now compressed 
adiabatically, that is, without addition or abstraction of heat. 
As a result of the compression the pressure and temperature 
increase; the compression pressure is given by the equation 



Vc 



&"- 



where r = — is the ratio of compression. The temperature 

is given by 

T 

L2 = r o.4 

T b 

If the atmospheric pressure, p b , is 14.7 lb. per square inch and 
the temperature t b is 60°F., or, T h = 460 + 60 = 520° abso- 
lute, then, with r = 4.8, p c = 9 X 14.7 = 132.3 lb. per square 
inch and T c = 1.873 X 520 = 973° absolute, or, t c = 973 - 
460 = 513°F. 

As a result of the explosion at c the heat of combustion is lib- 
erated and is used in heating up the charge. The heat of com- 
bustion may be assumed to be 80 B.t.u. per cubic foot of 
charge admitted, or 80 (v b — v a ) B.t.u. The weight, w } of gas 
in the cylinder is given by the perfect-gas equation 

144 pv = WRT, 

where R is the gas constant, which may be assumed to have the 
value 52 for the usual explosive mixture. For the conditions 
assumed, the weight of gas is 

144 pv 144 X 14.7 n n7Q . 

W = -RT~ = 52 X 520 * = Q ° 784 Vb ' 

The heat required to raise 1 lb. of this gas 1°F. while the volume 
remains unchanged is called the specific heat at constant volume, 
C v , and may be taken as 0.171 B.t.u. The rise in temperature 
during explosion is given by the equation 



ENGINE EFFICIENCIES AND CAPACITIES 



15 



Heat of explosion, H, 
temperature, or 



H 



weight of gas X specific heat X rise of 

= wC v {T d - T c ) 



For the conditions assumed, 

800& - v a ) = 0.0784 v b C v (T d 



To) 



and since 



^ = 4.8 

V a 



1 



o o 

Tc = 80 X 4^ X ^07 



4,740° 



and the explosion temperature, T d , — 4,740 + 973 = 5,713°. 
The explosion pressure, pa, is given by the equation 

Pd = Td 
p c T c 



OT,p d 



5,713 
973 



X 132.3 = 778 lb. per square inch. 



The expansion curve in the ideal cj^cle is adiabatic, so that the 
equations are the same as for the compression. The pressure 
p e at release or beginning of exhaust, e, is given by 
pa '« * 1 - i 

Ta 



and 



>d = /Ve\ 



— r 0.4 



.1.4 



778 
Consequently, p e = -q- = 86.4 lb. per square inch 



and 



T e 



5,713 



3,060° 



1.873 
The actual indicator card is shown 
in Fig. 11. At the end of the exhaust 
stroke the pressure in the cjdinder is 
above that of the atmosphere; usually, 
p a = 15.4 to 17 lb. per square inch. 
This pressure falls as the piston 
moves down and the fresh charge is 
drawn in. Mixing with the residual 
burned gas this charge is somewhat 
heated, the temperature, t b , being usu- 
ally from 170 to 260° and the pressure, p&, 12 to 14 lb. absolute. 
This pressure depends on the engine speed and on the resistance 



c -400 

I 300 

-".zoo 

in 
<v 

3 ioo 



Admission-' 

Vol umes 

Fig. 11. — Actual indicator 
card of Otto-cycle engine. 

























W- 










\ 










s 


^h^ 






c 




^ 




'Exhc 


X£ 


*»„. 




- — .e 


ri/5^*>^ci« 


?&S/orr 




A 



16 



THE AIRPLANE ENGINE 




encountered by the fresh charge in the carburetor, manifold, 
and inlet valve. 

The compression curve be is not adiabatic, but may be 
represented by the expression p c v c n = PbVb n , where n has a value 
between 1.23 and 1.35. The compression pressure, p c , is deter- 
mined by the clearance volume: 110 to 120 lb. gage is near the 
average of good practice for water-cooled engines. The com- 
pression pressure tends to fall with increase of engine speed in 

consequence of the increase 
of vacuum in the cylinder 
at the moment of closure of 
the inlet valve. The curve 
for the Liberty engine (Fig. 
12) shows this effect. The 
maximum temperature, t d , 
following ignition, is usu- 
ally from 2,500 to 3,200°F., and the maximum pressure, pa be- 
tween 300 and 400 lb. When p c is increased, p d also increases. 
The expansion curve de also is not adiabatic but may be rep- 
resented by PdVd n = p e v e n , in which n = 1.27 to 1.5; the low values 
are realized in cases where combustion continues after expansion 
has been well started. The terminal pressure, p e , is from 38 to 
75 lb., and t e from 1,200 to 2,000°F. The pressure during exhaust 
varies from 17 to 15.4 lb. per square inch. 

The compression ratio, -> is fixed by the clearance and may be 

varied in a given engine by the use of pistons with heads of dif- 

Table of Compression Pressures, Pounds per Square Inch 



1200 1400 1600 1800 2000 
Engine Revolu+ions per Minu+e 

Fig. 12. — Variation of compression pressure 
with engine speed. 





Compression ratios 




3.5 


4.0 


4.5 


5.0 


6.0 


7.0 


1.25 
1.30 
1.35 
1.41 


59.9 
63.7 
67.8 
73.1 


70.6 

75.8 
81.2 
88.3 


81.9 

88.3 

95.2 

104.2 


93.2 
101.3 
109.8 
120.9 


116.5 
128.4 
140.4 
156.4 


143.0 
156.9 
172.9 
194.3 



ferent shapes. It is the chief factor in determining p c . The 
accompanying table is based on pb = 12.5. A high compression 



ENGINE EFFICIENCIES AND CAPACITIES 17 

ratio increases power output and efficiency. It also increases 
the temperature at the end of compression since 

T c _ t c + 460 /PA*-' 

T b t b -f 460 \pj 
A high value of t c may cause preignition and thus fixes a limit of 
compression ratio which must not be exceeded. Usual values are 
from 4.5 in hydroplanes to 5.6 for high-altitude land machines. 
Values of compression ratio are given for various engines on pages 
66 to 70, where it is seen that high values result in increased 
power output per unit of cylinder volume. Values of six or higher 
may be used for high altitude work, but the engines will develop 
preignition if operated at full throttle near the ground. When 
operated on partial throttle the entering charge is of reduced 
weight, both because of lower pressure and because of dilution 
of the fresh mixture with a relatively greater weight of burnt gas 
remaining over from the previous cycle. Consequently the heat 
developed per cycle is less and the mean temperatures in the 
cylinders are reduced. 

Efficiency. — If the working substance in the cylinder followed 
the laws of a perfect gas 

pv = RT 
and C v = constant 

and if the combustion were instantaneous and complete, the 
efficiency of the cycle would be equal to 



- <r « 



where r is the compression ratio. It is here assumed further 
that the cycle takes place without any heat exchange between 
the working charge and the cylinder. The efficiency so found is 
the highest possible efficiency for an engine operating on the Otto 
cycle and could be attained only under the conditions stated 
above. It is sometimes called the air -cycle efficiency. Its value 
for various compression ratios is given in the following table : 

r=3.50 4.00 4.50 5.00 5.50 6.00 6.50 7.00 
e = 0.40 0.43 0.45 0.48 0.50 0.51 0.53 0.54 

The efficiencies actually obtained in airplane engines are seldom 
greater than 60 per cent of these values. For instance, with 
r = 5.5, the actual efficiency will not exceed 0.6 X 0.5 = 0.3 



18 THE AIRPLANE ENGINE 

and will in general be between 0.25 and 0.3. This considerable 
discrepancy between the actual performance of the engine and 
the air cycle efficiency is due to a variety of causes, the principal 
of which are as follows : 

I. The theoretical cycle assumes that the total heat (lower 
heat value) of combustion of the fuel taken into the cylinder is 
utilized during explosion in heating up the working mixture. 
This is not actually the case for two reasons : 

(a) The whole of the heat of combustion is not evolved during 
explosion because combustion is not instantaneous, so that 
combustion will continue for part (or with incorrect mixture, 
for the whole) of the expansion stroke, thereby reducing the 
amount of heat available for conversion into work. Furthermore, 
complete combustion at the end of explosion is not attainable 
because chemical equilibrium requires the presence of a certain 
amount of hydrogen and carbon monoxide. Their existence is 
commonly ascribed to dissociation. The amount of heat suppres- 
sion from these causes is not considerable in a high-grade engine 
operating with gasoline and with a good mixture. 

(b) Some of the heat actually evolved goes to the cylinder 
walls by radiation and conduction. The total heat so going 
to the walls in airplane engines is from 25 to 30 per cent of the 
total heat of combustion. If the heat were abstracted from the 
burning mixture during the explosion it would result in a loss 
of efficiency of the same magnitude. The actual passage of 
heat from the mixture to the walls continues from the middle 
of compression to the end of exhaust. Throughout the whole of 
this time the gases are hotter than the walls. The heat flows 
in the opposite direction during the admission period and the 
first part of the compression, but the amount of heat thus flowing 
is small, as the temperature difference between the walls and the 
gases is small. Such heat as passes off during the explosion 
and the first part of the expansion stroke may be regarded as 
entirely lost to the engine; the heat flow to the walls near the 
end of expansion and during exhaust is no loss at all, as it is 
necessary to discharge the hot gases and it is immaterial, from 
the point of view of efficiency, whether the heat is carried away 
by the jacket water or in the exhaust gases. 

General experience would indicate that more than one-half 
of the total heat given to the jacket may be regarded as abstracted 
by radiation or conduction from the working substance during 



ENGINE EFFICIENCIES AND CAPACITIES 19 

explosion and the early part of expansion. It should be noted 
in this connection that the heat of the jacket water includes 
most of the friction work between the piston and cylinder, which 
is a considerable fraction of the total friction of the engine. As 
the jacket heat is 25 to 30 per cent, the heat lost during explosion 
by radiation and conduction may be taken as not more than 12 
to 15 per cent of the heat of complete combustion. 

II. The working substance is not a perfect gas and, in particu- 
lar, it is not true that the specific heat at constant volume is a 
constant. It is found on investigation that the gases (C0 2 , N 2 , 
H2O etc.) which are present in the cylinder after explosion have 
specific heats which increase considerably with increase of tem- 
perature. These specific heats follow the equation 

C v = a + bt 

where a and b are constants. 

The efficiency of the cycle is diminished as a result of this 
increase of specific heat. The immediate result is that the rise 
of temperature during explosion, for a given amount of fuel 
burned, is diminished and consequently the pressure, pd (Fig. 
10), is lower than would be realized with constant specific heat. 
The expansion curve de is consequently lowered and the work 
of the cycle diminished. The efficiency with adiabatic expan- 
sion and compression but with variable specific heat is given 
by the expression, 1 



E ~ »{l - ±[(l - e)T d + T>] 



where e is the air-cycle efficiency and T d and T b are the abso- 
lute temperatures at the end of explosion and beginning of 
compression respectively. The constants a and b have values 
of about 0.194 and 0.051 X 10~ 3 for the average working mix- 
ture. For the conditions customarily met in airplane engines 

E . 

the ratio of the two efficiencies — is about 0.80; in other words, 

e ' ' 

the theoretical efficiency with the actual working substance is 
only 80 per cent of that which would be attainable if these sub- 
stances were perfect gases. 

The actual working substance consists almost exclusively of 
nitrogen, water vapor and carbon dioxide. All three of these 
substances show a considerable increase in specific heat with rise 

1 Wimperis, The Internal Combustion Engine, p. 85. 



20 



THE AIRPLANE ENGINE 



in temperature and the last two dissociate at high temperatures, 
especially at low pressures. The following tables 1 give mean 
specific heats at constant volume, and percentage dissociation. 

Mean Specific Heats at Constant Volume (in B.t.u. per degree 
Fahrenheit) between 200°F. and the Stated Temperature 



Temperature, degrees 
Fahrenheit 


930 


1,830 


2,730 


3,630 


4,530 


5,430 


Nitrogen 


0.185 
0.350 
0.187 


0.188 
0.385 
0.217 


0.196 
0.425 
0.229 


0.205 
0.468 
0.238 


0.214 
0.540 
0.247 


0.225 


Water vapor. 

Carbon dioxide 


0.623 
0.249 



Dissociation, Per Cent 







Pressure 


in atmospheres 




Temperature, 












degrees 


0.1 


1.0 




10 


100 


Fahrenheit 














H 2 


2,730 


0.043 


0.02 




0.009 


0.004 


3,630 


1.25 


0.58 




0.27 


0.125 


4,530 


8.84 


4.21 




1.98 


0.927 


5,430 


28.4 


14.4 




7.04 


3.33 




C0 2 


2,730 


0.104 


0.048 




0.0224 


0.01 


3,630 


4.35 


2.05 




0.96 


0.445 


4,530 


33.5 


17.6 




8.63 


4.09 


5,430 


77.1 


54.8 




32.2 


16.9 



Calculation of the theoretical efficiency, taking into account 
both the variable specific heats and dissociation, shows that this 
1 Tizard and Pye, The Automobile Engineer, Feb., 1921. 



ENGINE EFFICIENCIES AND CAPACITIES 



21 



efficiency, with a mixture giving maximum efficiency, is repre- 
sented very closely by the equation 



•-»-© 



(2) 



The heat loss to the walls reduces the actual efficiency below the 
theoretical values. With the very best design of cylinder and 
optimum operating conditions the highest attainable indicated 
thermal efficiency is given fairly accurately by the equation 



E 



- ■ - © 



(3) 



A comparison of these efficiency values is given in the following 
table. There are added the best results obtained by Bicardo 1 
on a special engine in which every known refinement was em- 
ployed with a view to raising the thermal efficiency. 





Cycle and 


Engine Efficiencies 






Efficiency 


Compression * 










ratio r 


Air cycle 


From equa- 
tion (2) 


From equa- 
tion (3) 


Ricardo's 

observed 

values 


4.0 


0.426 


0.336 


0.296 


0.277 


4.5 


0.452 


0.359 


0.314 


0.297 


5.0 


0.475 


0.378 


0.332 


0.316 


5.5 


0.494 


0.396 


0.348 


0.332 


6.0 


0.512 


0.411 


0.361 


0.346 


6.5 


0.527 


0.424 


0.375 


0.360 


7.0 


0.540 


0.437 


0.386 


0.372 


7.5 


0.553 


0.449 


0.396 


0.383 


8.0 


0.565 


0.460 


0.406 





The difference between the air-cycle efficiency (constant 
specific heat and no dissociation) and the theoretical efficiency 
of the cycle using imperfect gases, with the properties given in 
the preceding tables, diminishes as the explosion temperature 
diminishes. In the hmiting case in which there is no fuel in the 
charge and consequently no rise of temperature at explosion the 
two efficiencies become equal. The less the fuel in the charge, or, 

1 Proc. Royal Aeronautical Society, 1920. 



22 



THE AIRPLANE ENGINE 



0.6 



the weaker the mixture, the more nearly does the cycle efficiency 
approach the air-cycle efficiency. 

Calculations by Tizard and Pye show the cycle efficiency to 
vary with the mixture strength as in Fig. 13. The curve for 
correct mixture shows the efficiency when the air-fuel ratio is 

/1\ 0.258 

chemically correct; the equation to the curve is E = 1 — (-) 

The 20 per cent weak curve is 
calculated for 20 per cent ex- 
cess of air which is usually 
about the limit of explodibility ; 
the improved cycle efficiency 
in this case is verified by en- 
gine tests which generally show 
maximum indicated thermal 
efficiency with about 20 per 
cent excess of air. The 50 
per cent weak curve represents 
a condition which cannot be 
attained in the normal Otto- 
cycle engine as it gives a non- 
explosive mixture; it can be 
realized by an injection of the 
fuel into the compressed air as 
in the Diesel cycle, or by hav- 
ing a stratified charge in the 
cylinder with an explosive 



0.4 



0.3 



0.2 

















1 


,f 












**y 














-^ 




> 


& 


& 






y 


// 



































5 6 7 & 9 10 

Compression Ratio 

Fig. 13. — Calculated and observed 
thermal efficiencies with various 
strengths of mixture and compression mixture Surrounding the ig- 
ratios. ., T .-, ,. 

niter. In any case the reali- 
zation of the higher efficiency of the weak mixture will be attended 
by reduced engine capacity. 

Another point of importance is brought out by the curves of 
Fig. 13. The ratio of cycle efficiency to air-cycle efficiency in- 
creases with the ratio of compression; that is, we may expect 
to realize a larger percentage of the air-cycle efficiency as the 
compression ratio increases. The ratio of the efficiencies for a 
compression ratio of 4 is 0.685; with a compression ratio of 
10 it is 0.735. This improvement is shown also in actual 
engines. The ratio of observed efficiency (Fig. 13) to the air- 
cycle efficiency rises from 0.65 at a ratio of compression of 4 to 
0.685 at a ratio of compression of 7. 



ENGINE EFFICIENCIES AND CAPACITIES 23 

III. The theoretical indicator diagram is not realized for 
still another reason. The admission and exhaust of the charge 
are attended by frictional resistance to the passage of the gas 
through the carburetor, inlet manifold, inlet valve, and exhaust 
valve. Moreover, as the flow of the gas is at high velocity, there 
must be a pressure drop to bring about this flow; with an inlet 
velocity of 250 ft. per second, this would amount to about 0.6 
lb. per square inch. The frictional resistance and velocity head 
cause a lowering of the admission pressure, a, a raising of the 
exhaust pressure, and the forming of the "loop" (Fig. 11) at 
the bottom of the indicator diagram. This loop represents the 
negative pumping or fluid friction work which the engine has to 
perform. Engine tests indicate that the pumping work increases 
rather more rapidly than the square of the engine speed; the 
actual amount of the work depends on the dimensions and 
arrangement of the engine. At 1,000 r.p.m. it will probably 
average about 4 per cent of the indicated work of an aviation 
engine. This means in an engine with 120 lb. per square 
inch, brake m.e.p. that the mean height of the loop is 5 lb. per 
square inch at 1,000 r.p.m. The pumping loss is a function of 
the gas velocity in the manifolds and through the valve ports. 
Its magnitude will vary from about 2 lb. per square inch with a gas 
velocity of 100 ft. per second to 8 lb. per square inch with a gas 
velocity of 200 ft. per second. The indicated work may properly 
be considered as being only the positive loop of the indicator card; 
the suction-exhaust loop is one of the engine friction losses. 

Minor factors affecting the area of the indicator card are the 
rounding of the "toe" of the diagram which results from the 
opening of the exhaust valve before the end of the expansion 
stroke in order to facilitate exhaust, and the departure of the 
expansion and compression curves from the theoretical adiabatic 
curves. 

IV. It is not usually practicable to determine directly the 
work done by the working substance in the engine cylinder or 
the indicated work. In all tests the power measured is the 
useful work or brake horse power which is what remains of the 
indicated work after some of it has been used up in overcoming 
the friction of the engine and in driving the water and oil pumps. 

Indicated work — friction work = useful work 

Useful work , , , . , «, - 

T—p — t — t r = Mechanical efficiency 

Indicated work 



24 



THE AIRPLANE ENGINE 



Friction Losses. — The mechanical losses in an engine may be 
divided into two groups : 

1. The losses due to bearing friction and the driving of such 
auxiliaries as valve gears, oil and water pumps, magnetos, etc. 

2. Piston friction. 

Tests by Ricardo show that to overcome the first group a 
mean effective pressure of from 1.5 to 3 lb. per square inch is 
usually required — the lowest figure applying to a large multi- 
cylinder engine. The distribution of these losses is about as 
follows : 

Bearings . 75 to 1 . 00 lb. per square inch 

Valve gear . 75 to . 80 lb. per square inch 

Magnetos . 05 to . 01 lb. per square inch 

Oil pumps 0. 15 to 0.25 lb. per square inch 

Water pump. 0.30 to 0.50 lb. per square inch 

Total 2 . 00 to 2 . 65 lb. per square inch 

Piston friction is the largest item of loss; its magnitude prob- 
ably results from the fact that the motion is reciprocating and 




1000 



1400 1600 1800 2000 

Engine Revolutions per Min. 



2200 



Fig. 14. — Mechanical efficiencies of airplane engines. 

that the film of oil on the walls is more or less carbonized by the 
high temperatures and consequently has a high viscosity. The 
magnitude is probably about 7 lb. per square inch of piston area. 

The total loss from bearing friction, piston friction and fluid 
friction in the best ungeared engines is from about 10.5 to 14 lb. 
per square inch of piston area, the lower figure referring to radial 
air-cooled engines and the higher to water-cooled engines. 
Taking 120 lb. per square inch as the brake m.e.p., these values 
correspond to mechanical efficiencies of 92.0 and 89.5. Tests 
(Fig. 14) indicate that friction work increases more rapidly than 
the engine speed but not so rapidly as the square of the speed. 

Taking into account the losses enumerated, it is possible to 
arrive at a fair approximation to the actual efficiency of an air- 
plane engine. Consider for example a high-grade engine with 



ENGINE EFFICIENCIES AND CAPACITIES 25 

a compression ratio of 5.5, using gasoline as fuel. For every 
100 B.t.u. (lower heat value) of heat of combustion we may 
expect a heat suppression (Item La) of 4 B.t.u., leaving 96 B.t.u. 
developed. Of this quantity, 13 B.t.u. will go to the walls by 
radiation and conduction (Item 1.6) before it can be utilized, 
leaving 83 B.t.u. The theoretical efficiency for a compression 
ratio of 5.5 is 0.396 (equation (2) p. 21). The theoretical work 
of the cycle is 0.396 X 83 = 32.9 B.t.u. The actual indicated 
work is thus 32.9 per cent of the heat of perfect combustion of 
the fuel and this quantity is usually spoken of as the indicated 
thermal efficiency, or the thermodynamic efficiency, E t , of the 
engine. It measures the efficiency of the engine in converting 
heat into work. 

As previously stated, this indicated work is not readily meas- 
urable. The useful or brake work may be taken as 85 per cent 
(Item IV) of the indicated work, or in this case, 0.85 X 32.9 = 
28.0 B.t.u. The thermal efficiency referred to b.h.p. is then 
28.0 per cent. This quantity may be compared with the results 
of tests on a high-grade engine. Such tests may be expected 
to show the consumption of about 0.50 lb. of gasoline per brake 
horse power hour. The fuel has a lower heat value of almost 
18,500 B.t.u. per pound. The thermal efficiency referred to 

, . Work of 1 b.h.p. hour, B.t.u. 2,545 

b.n.p. is Heat of combustionof the fuel, B .t.u. ~~ 0.50 X 18,500 

= 0.275, which agrees very closely with the calculated efficiency. 
A reduction in any of the itemized losses will increase the final 
efficiency. 

Mean Effective Pressures. — The mean effective pressure 
(m.e.p.) of a gas engine is that gas pressure on the piston which, 
if maintained constant for one stroke of the engine, would do as 
much work as is actually done in the two revolutions of the cycle. 

In aviation engines the m.e.p. is practically always obtained 
from the brake horse power and is called the brake m.e.p. It is 
given by the equation 

brake m . e .p. = ^- X 33 000 '_ »•**• 

where d is the cylinder diameter in inches, s is the stroke in 
inches, N is the revolutions per minute, and n is the number of 
cylinders. 



26 THE AIRPLANE ENGINE 

The brake mean effective pressures usually given are computed 
from the b.h.p.; values range from 70 to 135 lb. per square inch. 
The true m.e.p. in the cylinder is this value divided by the mechani- 
cal efficiency, E m . 

Torque and Power. — If p = actual brake m.e.p., the average 
useful force exerted in the cylinders of a four-cycle engine is 

IT V 

jnd 2 X t = 0.1964 pnd 2 lb. This force is maintained while the 

piston moves during each revolution 2s in., or s -s- 6 ft. and the 

work done in foot-pounds is 0.1964 pnd 2 X ~ = 0.03273 pnd 2 s. 

Torque is the average turning moment and is numerically equal 
to the force continuously exerted at the propeller at 1 ft. radius. 
This is exerted, during each revolution, over a distance of 2w = 
6.2832 ft. The work being equal to that already computed, 
the torque in pound-feet is Q = 0.03273 pnd 2 s +- 6.2832 = pnd 2 s 
+ 192. 

For s = 7, d = 5, n = 12, p = 120; Q = 1,315 lb-ft. The 
actual torque varies, but has this average value. 

If S — piston speed, feet per minute = -^z-i the b.h.p. is H B — 

\&\ Sn -T- 33,000 = pd 2 Sn 4- 168,000 = Q X N -5- 5,250. Thus 

for 1,600 r.p.m., in the preceding example, H B = (1,315 X 
1,600) 4- 5,250 = 401. 

Capacity and Volumetric Efficiency. — The weight of fuel 
mixture taken into the ideal engine (Fig. 10) is given by the gas 
equation 

144p (v b — v a ) 



Wn = 



R T 7 



TV 

where Vb — v a = a d 2 s is the volume of the mixture admitted and 

T m is the absolute temperature of the external air. 

Actual engines do not draw in weights equal to that expressed 
by the above equation. The weight of mixture actually admitted 
is the difference between the weight present at the points b and 

a, Fig. 10. The weight present at any point is w = ~pm~' 

, , +u • ,, A ... -, • 144p 6 » fc lUp a V a 

therefore the weight admitted is r> m — Dm — 

til b HI a 

The ratio of the weight actually admitted to that which would be 



ENGINE EFFICIENCIES AND CAPACITIES 



27 



admitted to the ideal engine is called the volumetric efficiency, E v . 

144 X 14.7 (v b - v a ) 



*-5-(^fS:-^i© + 



RT, 



Vb 



Writing — = r, the above reduces to 

rr Tm (PbT _ Va\ 

* v 14.7(r -l)\T b TJ 

Taking t m = 100, r = 5, p b = 12, fe = 200, p a = 16, * a = 900 
we find E t = 0.76. 

The volumetric efficiency is determined mainly by two factors, 
the temperature, T b , and the pressure, p b , at the end of admission. 



,3.3 



£3. 



.2. 
o 

52/ 



v 



-^% -K- 



2-3 200 400 600 800 1000 1200 1400 1600 

R.P. M. 

Fig. 15. — Volumetric efficiencies of hot and cold engines. 



The temperature of the mixture rises during admission as a result 
of the addition of heat from the hot interior surfaces of the cylin- 
der. Comparative tests of the volumetric capacity of an engine 
(1) when being motored over cold, and (2) in ordinary operation, 
show that the heating effect decreases slightly as the speed of the 
engine increases in consequence of the shorter time available for 
the transmission of heat. The pressure drop, however, increases 
continuously with increase of engine speed. Figure 15 1 gives 
curves of weight of charge taken into the cylinder per revolution 
for an engine of high valve resistance. The effect of the tempera- 
ture rise in reducing the volumetric efficiency is from 12 to 15 per 
cent and would be appreciably greater but for the evaporation of 
the fuel, which, through the abstraction of the latent heat of evap- 
1 Judge, High-speed Internal-combustion Engine, p. 161. 



28 



THE AIRPLANE ENGINE 



oration, reduces the rise of temperature of the charge by 30 to 
40°F. The variation of volumetric efficiency of the Liberty-12 
with engine speed is shown in Fig. 16. In the same figure there 
is also shown the volumetric efficiency of the Hispano-Suiza 300 
engine, but in this case the volumetric efficiency is given as the 
ratio of the weight of air actually admitted to the weight of a 
volume of air equal to piston displacement and of the density 
of the air in the inlet manifold. Measured in this way the volu- 
metric efficiency is 95 per cent at 1,600 r.p.m.; this corresponds 
to 93 per cent when compared with air at room density. The 
broken line shows the volumetric efficiencies compared with air 
at room density. 



no 



31oo 



"!F 90 



80 



70 































Hisf. 


?C*r>0 


r^^i 


:r 




- — 


L/bet 












-*/ 




*^-^ 


■ — . 






c 


r ~~~- 




















~^T 





■§ " 1200 1400 1600 1800 2000 

R.RM. 

Fig. 16. — Volumetric efficiencies of airplane engines. 



The pressure drop during admission is much more variable in 
different engines than is the temperature rise. Its magnitude 
depends on the pressure drop through the carburetor and the 
size and arrangement of the manifolds and the inlet valve. In 
every case it will increase rapidly with increased engine speed; 
its actual magnitude will usually be small for low speeds, 

The pressure drops in the manifolds of two engines are given 
in Fig. 17. It will be seen that with wide-open throttle the 
pressure drop is nearly proportional to the engine speed. With an 
engine loaded with a propeller, change of speed is obtained only 
by varying the opening of the throttle valve; the manifold 
vacuum increases rapidly as the throttle valve is closed. The 
pressure in the cylinder is considerably less than that in the 
manifold because of the valve resistances. 

The volumetric efficiency in engines of good design will be 
from 80 to 85 per cent. With low speeds and other exceptionally 
favorable conditions, values as high as 90 to 92 per cent have 
been recorded. 

The maximum possible volume of charge admitted per 



ENGINE EFFICIENCIES AND CAPACITIES 



29 



cycle in the ordinary engine is the volume enclosed at the instant 
of valve closure less the clearance volume. The admission 
valve always closes past the dead center. If the closing angle 
is 45 deg. late the piston will have returned about 12 per cent of 
its stroke and the maximum possible volumetric efficiency will 
be 88 per cent. Occasionally the operating conditions and induc- 
tion pipe length may be such as to give more than atmospheric 
pressure in the cylinder at the instant of closure which would 
result in increased volumetric efficiency. 




1400 1600 1800 2000 2200 2400 

Revolutions per Minu+e of Engine 

Fig. 17. — Intake manifold depressions with full throttle and with propeller 

load. 



It should be noted that the capacity (as affected by volumetric 
efficiency) and thermal efficiency of an engine are not necessarily 
related to one another. The diminution in capacity of an engine 
resulting from heating of the entering charge, from a high 
carburetor or inlet-valve resistance, or from diminishing speed, 
may or may not result in a change in efficiency, and the change, 
if it takes place, may be either an increase or a decrease. A 
given engine may at one time be developing 200 h.p.; at another 
time 250 h.p. ; the efficiency may, however, be the same in both 
cases, although it tends to be lower for the lower h.p. because of 
the approximate constancy of engine friction, which makes the 
efficiency referred to the b.h.p. less at light loads. 

Units of Capacity. — In determining the size of a projected 
engine, or in comparing the performance of existing engines, it 
is desirable to have some standard unit for measuring the specific 
capacity. The most common unit is the piston displacement in 



30 THE AIRPLANE ENGINE 

cubic inches per brake horse power, or the brake horse power 
developed per cubic foot of piston displacement. The piston 
displacement is the displacement per stroke of one cylinder 
multiplied by the number of cylinders. As the horse p'ower 
varies almost directly as the engine speed, the above units do 
not really lead to a satisfactory comparison of engines operat- 
ing at different speeds. For this purpose it is better to state 
the capacity at 1,000 r.p.m., deducing this capacity from the 
actual performance by the use of the assumption that horse 
power is proportional to engine speed. For example, the Lib- 
erty-12 engine, 5 by 7 in., develops 400 h.p. at 1,700 r.p.m. The 

7T 

piston displacement per cylinder is j X 5 2 X 7 = 137.4 cu. in. 

per stroke; the total piston displacement is 12 X 137.4 = 1,648.8 
cu. in.; the piston displacement per brake horse power is 1,648.8 
■f- 400 = 4.12 cu. in. The brake horse power per cubic foot of 
piston displacement is 12 3 ^4.12 = 420 h.p.; the piston dis- 
placement per brake horse power at 1,000 r.p.m. is 4.12 X 

■/Ann = 7.0 cu. in.; the b.h.p. per cubic foot of piston displace- 
ment at 1,000 r.p.m. is 420 X y^ = 247. 

The Hispano-Suiza engine, with 718.9 cu. in. displacement, 
develops 150 h.p. at 1,450 r.p.m. This corresponds to 4.8 cu. 
in. per horse power; or 6.96 cu. in. per horse power at 1,000 r.p.m. 
This last figure shows that the Hispano-Suiza and Liberty engines 
have practically the same capacities per cubic inch of piston 
displacement per minute. 

The fixed-cylinder radial air-cooled ABC Dragonfly nine- 
cylinder engine, with 1,389.3 cu. in. displacement, develops 310 
h.p. at 1,650 r.p.m. This corresponds to 4.48 cu. in. per horse 
power, or 7.38 cu. in. per horse power at 1,000 r.p.m. 

The lower specific capacity of rotating-cylinder engines is 
illustrated by the nine-cylinder Gnome, with a piston displace- 
ment of 770 cu. in., which develops 104 b.h.p. at 1,200 r.p.m. 
This corresponds to 7.41 cu. in. per horse power, or 8.89 cu. in. 
per horse power at 1,000 r.p.m. 

The above figures may be regarded as characteristic of the 
different types. 

Tests of Performance. — The results obtained on the test of an 
engine will vary greatly with a number of factors such as the air 



ENGINE EFFICIENCIES AND CAPACITIES 



31 



pressure and temperature, kind of fuel, type and dimensions of 
carburetor, temperature of jacket water and of lubricating oil, 
and condition of engine. For example the Liberty 12 has shown 
a brake horse power at 1700 r.p.m. which varies from 380 to 480 
b.h.p. 

The following tests are reported under sea-level conditions: 



Engine 


Revolu- 
tions 
per 
minute 


Brake 
m.e.p., 
lb. per 
sq. in. 


Brake 
horse 
power 


Friction 
horse 
power 


Me- 
chanical 
efficiency 


Hispano-Suiza-180 


1,200 
1,500 
1,700 
1,900 


114.0 
119.1 
117.7 
111.0 


121.8 
159.0 
178.0 
187.0 


16.8 
23.5 

28.4 
34.2 


0.88 
0.87 
0.86 
0.84 


Liberty-12 


1,200 
1,400 
1,600 
1,800 
2,000 


118.0 
119.5 
119.5 
117.6 
104.0 


295.0 
348.0 
398.0 
442.0 
433.0 


27.6 
38.3 
49.1 
65.4 
88.0 


0.91 




0.90 
0.89 
0.87 
0.83 



A plotting of test results on the Liberty 12 is shown in Fig. 18. 
These figures illustrate the usual laws of performance. The 



500 
130 
460 
440 
420 
|400 
£-380 

o360 

x 
340 

320 

300 

280 

260 



























Me 


"A E 


ma 


'ncu 




/ 








90 
80 










































Therm a 


/ , 


'Efficiency „ 




^ 


> 








/ 






y 










t 


/ 




/ 














V 


s 


7 














) 




y 
















< 


/ 


/ 




















/ 

































































1200 1400 1600 1800 2000 

R.R M. 

Fig. 18. — Performance curves of Liberty-12 engine. 



mean effective pressure, p, and consequently the torque, reach 
maximum values at some moderate speed. The power increases 
with increasing speed, but at a rate which diminishes after the 



32 THE AIRPLANE ENGINE 

maximum value of p has been reached. If p were constant the 
power would vary directly with the speed. Cylinder cooling 
reduces p at low speeds; high resistance through ports and 
passages reduces volumetric efficiency and p at high speeds (Fig. 
16). Maximum power is reached when the rate of decrease of p 
with engine speed is equal to the rate of increase of engine speed. 

Figure 79 gives results of trials on a 230 h.p., six-cylinder Benz 
engine. Here the failure of the power to increase proportionately 
to speed is clearly shown. Maximum mean effective pressure 
occurs at 1,050 r.p.m. and maximum power at 1,650 r.p.m. The 
fuel consumption rate is also shown. The economy is practically 
constant over the speed range 900 to 1,200 r.p.m. 

It will be observed that throttling the engine increases the 
fuel consumption per horse power. The full-line curves show 
the performance when the throttle is wide open and the engine 
is loaded until it assumes the desired speed. The broken lines 
show the performance when the engine had the propeller load 
only; in this case the engine speed can be reduced below 1,400 
r.p.m. only by partly closing the throttle; speeds above 1,400 
r.p.m. are not possible with the propeller load. 

Figure 143 is for a rotary-cylinder engine. In this case the 
effective horse power is the indicated horse power minus the engine 
friction loss and minus the windage loss; the rapid increase of 
windage loss when the engine speed increases makes the net 
horse power a maximum at about 1,250 r.p.m. 

Correction to Standard Atmospheric Conditions. — The pub- 
lished results of engine tests may give either the actual horse 
powers observed or these horse powers corrected to some standard 
atmospheric condition. The latter is much preferable as it will 
permit an immediate comparison of engine performances. The 
best measure of capacity is the brake m.e.p. That engine which 
has the maximum m.e.p. is developing a horse power on the 
smallest piston displacement. But to make such direct com- 
parison the operating conditions must be the same or else the 
results must be corrected to allow for differences. Such correc- 
tions can be readily applied to differences in atmospheric pressure 
and temperature. It has been commonly assumed that the 
horse power developed is proportional to the density of the 
atmosphere. The density is proportional to the atmospheric 
pressure, and inversely proportional to the absolute temperature. 
If the standard conditions are 14.7 lb. pressure (29.92 in. of 



ENGINE EFFICIENCIES AND CAPACITIES 



33 



mercury) and 32°F., and the observed horse power is P at 
pressure p and temperature t, the corrected horse power is 

M7 460 + * 
rc ~ r V 492 

In most of the published tests the correction to standard 
conditions has been made by use of this equation. Tests at the 
Bureau of Standards indicate, however, that the temperature 
correction in this equation is excessive and that more accurate 
results are obtained from the equation 

_ 147 920 + * 
c V 952 

The m.e.p. is corrected in exactly the same manner. 

There is no method for directly comparing two engines which 
are using different fuels or mixtures of different strengths 

Influence of Strength of Mixture on Capacity and Efficiency. — 
The effect of strength of mixture has been investigated by 
Berry 1 on automobile engines; his results are supported by 
the investigation of others on 
both automobile and aviation 
engines (see p. 260). For any 
-constant engine speed and con- 
stant throttle opening, they 
show (Fig. 19) that the max- 
imum power is obtained with a 
comparatively rich mixture, 
and that for maximum effici- 
ency a weaker mixture must 
be used. As the throttle is 
closed the mixture for maxi- 
mum efficiency (Fig. 21) be- 
comes richer and at the lowest 
loads coincides with that for 
maximum power. The speed 
of the engine has no appreci- 
able influence on the variation of engine power with strength of 
mixture (Fig. 22). Maximum power (Fig. 20) is obtained with 
0.08 lb. of gasoline per pound of air, or 12}^ lb. of air per pound 
of gasoline. Maximum efficiency is obtained with a mixture 
of 15 to 16 lb. of air per pound of gasoline at full load, but this 

1 Trans. Am. Soc. Mech. Eng., 1919. 























N5% 










Pow 


er ^ 


h 










'1 
















£ 












\ 




6 






































LEANER 






RICHER 







v 0.05 0.06 0.07 0.08 0.09 0.10 0.1 1 0.12 0.13 
Pounds of Gasoline per Pound of Air in Mixture 

Fig. 19. — Variation of power and 
thermal efficiency with strength of mix- 
ture, at full throttle. 



34 



THE AIRPLANE ENGINE 



mixture must be made richer as the load diminishes and becomes 
123^2 lb. of air per pound of gasoline at lowest loads. 




0.05 0.06 0.07 0.08 0.09 0.10 0.11 0.12 0.13 
Pounds of Gasoline per Pound of Air in Mixture 

' Fig. 20. Fig. 21. 

Fig. 20. — Variation of engine power with strength of mixture at constant engine 

torque and varying speed. 
Fig. 21. — Variation of thermal efficiency with strength of mixture at constant 
engine torque and varying speed. 





















&£ 


me . 


*>Peec 


r^ 








/ 










/ 


/ 


P° 






"■■*>, 


w 


j 




§ 
g 








s 


A 






^ 


'£oo 


\ 


I 




1 








* 


\ 


$ 






—~2Po 


\ 






L- 






■—222 




J 

LEAh 


IER 


t 












i 












i 




RICHER 







0.05 006 0.07 Q08 0.09 0.10 0.11 0.12 0.13 
Pounds of Gasoline per found of Air in Mixture 



Fig. 22. 



30 



=r>26 



fc 22 



















^v^ 


k 




^v 


% 












\2- 

















8 16 14 12 10 

Ratio of Air to Fuel by Weight 

Fig. 23. 



Fig. 22.- 
Fig. 23.- 



-Variation of engine power with strength of mixture at constant engine 

speed and varying torque. 
-Maximum thermal efficiencies of certain fuels with varying strength 
of mixture. 



Tests by Watson 1 on an automobile engine with gasoline, 
benzol, and wood alcohol as fuel show (Fig. 23) the variation in 
1 Proc. Inst. Aut. Eng., 1914. 



ENGINE EFFICIENCIES AND CAPACITIES 



35 



efficiency of these fuels with strength of mixture. The com- 
parison given by these curves is not, however, complete, since 
the same compression ratio was used for all three fuels. With 
alcohol it is possible to increase the compression pressure con- 
siderably without danger of preignition and without producing 
excessive explosive pressures; with benzol the compression ratio 
can similarly be increased slightly. The efficiency with alcohol 
could probably be raised to at least 35 per cent by the use of a 
higher compression ratio. 

Influence of Air Temperature on Capacity and Efficiency. — 
As already pointed out in the discussion of volumetric efficiency 
(p. 27) the temperature of the air admitted to the cylinder has 



214 
o 

S-12 



C 6 

t- 

* 4 
Z 
Q 







1 










1 

-5 


ficTemp 


Deg. 


£?. 






' 




) 




rT--- 






>p\ 




& 




-- 


~zrz 


3 


^> T 


Or 


i 


k 








y& 


$& 


y 


1 
-1 

6 


















































LEANER 






RICHER 






105 


06 ( 


.07 0. 


38 0. 


39 0. 





1 


2 0.1 




0.05 0.06 0.07 0.08 0.09 0.10 0.11 0.12 0.13 
Pounds of Gasolene per Pound of Air in Mixture 



Fig. 24. Fig. 25. 

Fig. 24. — Variation of engine power with strength of mixture at various air 

temperatures and full throttle. 
Fig. 25. — Variation of thermal efficiency with strength of mixture at various 

air temperatures and full throttle. 



a considerable influence on the power developed. The tests 
at the Bureau of Standards show the diminution in power with 
increase in temperature to be proportional to half the increase 
in absolute temperatures, for the range of temperature from 
4 to 120°F. For example, if the absolute temperature of the 
air increases from 500 to 580°, or 16 per cent, the power of the 
engine will decrease 8 per cent. Berry's tests on automobile 
engines (Figs. 24 and 25) show that this law does not hold for a 
higher temperature range. The decrease in power is roughly 
proportional to one-third the increase in temperature (Fig. 24). 
The efficiency (Fig. 25) is also seen to diminish with increase of 
air temperature but through a much smaller range. 

Tests by Berry with air at a temperature lower than 80°F. 



36 



THE AIRPLANE ENGINE 



showed a rapid falling off in capacity and efficiency. These 
tests, however, were carried out with a commercial gasoline 
of low volatility as compared with the gasolines specified for 
airplane engines. In all cases maximum power is obtained with 
the lowest air temperature which will permit satisfactory dis- 
tribution and vaporization of the fuel. This temperature 
depends not only on the volatility of the fuel but also upon the 
manifold design. A" hot spot " between the carburetor and mani- 
fold (heated by the exhaust gases), on which the liquid spray 
from the carburetor impinges, causes vaporization of part of the 
fuel without heating up the air appreciably and is found to 
result in a better distribution of the mixture to the different 
cylinders and in improved engine operation when the air supply 
is cold. With this device it is possible to lower the temperature 
range of the entering air somewhat without a falling off in capacity 
or efficiency. 

Influence of Throttling on Efficiency. — It is generally found 
that thermal efficiency tends to increase as the power is cut down 



0.65 



C 0.60 



0.55 



0.50 



0.45 



One half load 




^'^ fhlT -^ 



1400 



1600 



1800 



Fig. 26. 



800 1000 1200 I4UU I6UU Ibuu 

, — Variation of specific fuel consumption with engine speed at various 
throttle positions. 



from maximum to three-fourths load. This phenomenon is prob- 
ably due mainly to improvement in the mixture at partial load. 
The carburetor is set to give maximum power for full throttle, 
which, as just shown, is obtained with an over-rich and conse- 
quently inefficient mixture. If the mixture becomes leaner at 
partial throttle, the economy will improve. What actually 
happens will depend primarily on the characteristics of the 
carburetor used. 



ENGINE EFFICIENCIES AND CAPACITIES 



37 



In Fig. 26 are given the fuel consumptions per brake-horse- 
power hour of a Liberty 12, at full, three-fourths and one-half 
loads. It will be seen that the efficiencies at full and three- 
fourths loads are substantially the same, but that there is a 
marked falling off at half load. The efficiency is again seen to 
increase with engine speed. 

Influence of Compression Ratio on Capacity. — Tests of a 150- 
h.p. Hispano-Suiza engine at 1,500 r.p.m. with various compres- 
sion ratios show the maximum attainable brake horse power to 
have been as follows: 



Ratio of compression 4.7 5.3 6.2 

Maximum brake horse power . . 160 . 165 . 169 . 



The percentage increase of power with increase of compression 
ratio is about the same as in the tests of a Liberty 12 engine, 
which at 1,600 r.p.m. gives the following results: 

Ratio of compression 4.9 5.5 

Maximum brake horse power 380. 398. 



The Influence of Revolutions per Minute on Capacity is shown 
in all performance curves (see Figs. 48, 50, 53, etc.). The fall in 



.£'40 

& 

^130 
o 

£l20 

CD 100 











Liber 


t y e } 








1 1 1 1 

H/spano - Sui zcr , ZOO 




Pack 


ardl 


f.f^_ 


^> 


e«=.r: 





~^. 


^CT 







'*-"•> 










"" 


HiSf 


jcrno 


-5u/'z 


~aZ~~ 


■^C 




^T 




ir- 


^ 














L 


iberfi 


//* 


^ 




'Pcrc 


karct 


IV 





























<*> 1000 1200 1400 1600 1800 2000 2200 2400 
Revolutions per Minute of Engine 

Fig. 27.- — Variation of mean effective pressure with engine speed for airplane 

engines. 

brake m.e.p. (Fig. 27) causes the b.h.p. to go through a maxi- 
mum; the efficiency practically always increases with speed, but 
becomes a maximum before the engine reaches maximum horse 
power. 

Influence of Jacket-water Temperature on Capacity. — Figure 
28 shows the effect on the capacity of a Liberty 12 engine of 
varying the temperature of the jacket water. The amount of 
water circulated was constant at any given speed but the inlet 
temperature was varied, thereby giving the series of outlet 



38 



THE AIRPLANE ENGINE 



temperatures indicated. It will be seen that the power increases 
as the cooling-water temperature decreases to about 100°F. At 
200°F. and 1,800 r.p.m. the power is only 417 h.p. while at 90° 
it is 436 h.p. 



400 



300 



280 





Correcte 


lh2S 


92Inc 


hes of Mercury ^ 


















rV 
















/ v / 


kV 
















W/ 














A 


v/ 
















w 


i 














4 




200°F 












m 




^I10°F 
~I50°F. 
-I30°E 
-JICF 














w 














/ 




"^~ 


-90°F 












M 
















A 


w 
















ffl 


/ 
















W 


















V 

































13 14 15 16 

R. P. M. Hundreds 



Fig. 28. — Variation of engine power with speed for various jacket-water dis- 
charge temperatures. 

Various tests have demonstrated that the friction horse power 
decreases with increase of jacket-water temperature, so that the 
above increase of brake horse power with lower jacket- water 
temperature must have resulted from a still greater increase in in- 

40 

30 

\r\esiwuui 

g3 0[— 1 | 1 -- 

x 

^20 

4- 

jio 











t .. 








Exha 




- 






















Res/ 


dual 




























B. 


H.P. 






, 




















Jack 


e-h 


























1200 1400 1600 1800 2000 

R.R M. 

Fig. 29. — Typical heat balance of airplane engine. 



dicated horse power. The increase is primarily due to improve- 
ment in volumetric efficiency resulting from less heating up of the 
charge as it enters. This phenomenon is shown in Fig. 15. 

Heat Balance. — Of the total heat of combustion of the fuel 
admitted to the engine cylinder, part is converted into brake 



ENGINE EFFICIENCIES AND CAPACITIES 39 

horse power, part goes to the water jacket, part escapes as heat 
in the exhaust gases, and the rest is lost in various ways such as 
incomplete combustion and radiation and conduction from the 
engine. A typical heat balance for the Liberty 12 is shown in 
Fig. 29 in which the percentages are in terms of the higher heat 
value of the fuel. It will be observed that the heat distribution 
does not change notably with engine speed. 



CHAPTER III 



ENGINE DYNAMICS 



Turning Moment. — The pressure on any single piston of a 
four-cycle engine is varying continuously throughout the cycle. 



500 



400 



300 



eo.o 



a- ioo 





































A 


































\ 
































\ 


































\ 
































s 




























































\, 































































































1 2 3 4 5 6 7 

Piston Travel, In. 
Fig. 30. — Indicator card of the Liberty-12 engine. 

If the indicator card is as in Fig. 30, the resulting total gas pressure 
on the piston of a 5-in. diameter, 7-in. stroke engine for various 

9000 
8000/- 



7000 

6000 

5000 

4000 

3000 

2000 

1000 



-1000 

-2000 

-3000 



rt 


A 






A- 


! > 1 1 I 1 
r otal Gas Pressure on Piston 










f 








8- Inertia Forces at 1700 Rpm. 










rf 








Z-Rzsultant Pressure along Cyli 


hderA 


x/'s 




°\ 


I 






























\ 


\ 
































V 


































i*** 


























A^ 






\ 


"> 






















i 


/ 






/ 


"*> 




















% 


< 




B 


>/ 


























\ 


S' 


/ 


/ 


























Bj 




S 






























V 



Fig. 31.- 



] 40 80 120 160 200 240 280 320 360400 440 480 520 9M WO 640 680 720 
Crank Angle Degrees 

Forces acting on the crank pin of the Liberty-12 engine. 



successive crank positions is represented by the curve A, Fig. 
31. The pressure transmitted to the crankpin is modified, 
however, by the inertia of the reciprocating masses of the piston 

40 



ENGINE DYNAMICS 41 

and connecting rod. During the first part of each stroke these 
masses are being accelerated; during the second part they are 
retarded. Hence the net useful force acting on the crankpin 
in a direction parallel to the cylinder axis is alternately less or 
greater than that shown by the curve A . 

Let W = Weight of reciprocating parts, pounds (complete 

piston and half of rod). 
n = Revolutions per minute. 
a = Angle turned through by crank, starting from its 

uppermost position, degrees. 
r = Crank radius, feet (half the stroke of the engine). 
I = Length of connecting rod (center to center of pins), 

feet. 

Then the accelerating force at any moment, in pounds, is 

v 
P a = 0.00034 Wn 2 r (cos a ± -, cos 2a), approximately, the + 

sign being used for the down stroke and the — sign for the up 

T 

stroke. The quantity (cos a + -cos 2a) is an approximation; a 

V 

y 2 cos 2a + sin 4 a 
more correct expression is cos a ± Values of 

(-2 - sin 2a ) 

this quantity for the range of ratios of I to r common in airplane 
engines are given in Table 1. The minus sign indicates negative 
acceleration from deg. to 180 deg., and positive acceleration 
from 180 deg. to 360 deg. Calculations made for the Liberty-12 
engine give the results shown in Fig. 31. The indicator card 
(Fig. 30) is plotted for 18 per cent clearance and a brake m.e.p. 
of 123 lb. per square inch, the exponents of the compression 
and expansion curves being taken as 1.32. The cylinder is 5 
by 7 in., the connecting rod 12 in. long and the weights of 
reciprocating parts are: piston complete with pin, 4.838 lb.; 
upper half of connecting rod, 1.225 lb.; total, 6.063 lb. The 
engine is assumed to make 1,700 r.p.m. The inertia forces, P a , 
calculated from the preceding equation, are plotted as curve 
B, Fig. 31. The algebraic sum of gas pressures A and inertia 
pressures B is shown by the resultant pressure curve C. 



42 



THE AIRPLANE ENGINE 



Table 1. — Inertia Factors 
I 2 



cos a + 



cos 2a + sin 4 a 



(- 2 -sin 2 a) 



Crank 


I 


I 


I 


I 


I 


Crank 


angle, 


- = 4 


- = 3.75 


- = 3.5 


- = 3.25 


- = 3.0 


angle, 


degrees 


r 


r 


r 


r 


r 


degrees 





1 . 2500 


1.2667 


1.2857 


1.3077 


1.3333 


360 


5 


1 . 2426 


1 . 2590 


1.2778 


1 . 2995 


1 . 3249 


355 


10 


1 . 2204 


1.2362 


1 . 2543 


1 . 2752 


1.2997 


350 


15 


1 . 1839 


1.1986 


1.2155 


1.2351 


1 . 2580 


345 


20 


1.1335 


1 . 1468 


1.1621 


1 . 1798 


1 . 2006 


340 


25 


1 . 0702 


1.0817 


1.0948 


1.1102 


1.1283 


335 


30 


0.9950 


1.0042 


1.0149 


1 . 0274 


1 . 0423 


330 


35 


0.9091 


0.9158 


0.9236 


0.9328 


0.9440 


325 


40 


0.8140 


0.8179 


0.8225 


0.8281 


0.8349 


320 


45 


0.7112 


0.7121 


0.7133 


0.7149 


0.7172 


315 


50 


0.6026 


0.6004 


0.5980 


0.5955 


0.5929 


310 


55 


0.4899 


0.4846 


0.4787 


0.4719 


0.4643 


305 


60 


0.3751 


0.3668 


0.3573 


0.3465 


0.3338 


300 


65 


0.2601 


0.2490 


0.2363 


0.2215 


0.2041 


295 


70 


0.1468 


0.1332 


0.1175 


0.0992 


0.0776 


290 


75 


0.0368 


0.0211 


0.0030 


-0.0182 


-0.0434 


285 


80 


-0.0682 


-0.0854 


-0.1055 


-0.1288 


-0.1567 


280 


85 


-0.1669 


-0.1851 


-0.2062 


-0.2309 


-0.2605 


275 


90 


-0.2582 


-0.2767 


-0.2981 


-0.3234 


-0.3536 


270 


95 


-0.3412 


-0.3594 


-0.3805 


-0.4052 


-0.4348 


265 


100 


-0.4155 


-0.4327 


-0.4528 


-0.4761 


-0.5040 


260 


105 


-0.4809 


-0.4965 


-0.5146 


-0.5358 


-0.5610 


255 


110 


-0.5373 


-0.5509 


-0.5665 


-0.5848 


-0.6064 


250 


115 


-0.5851 


-0.5962 


-0.6090 


-0.6237 


-0.6411 


245 


120 


-0.6249 


-0.6332 


-0.6427 


-0.6535 


-0.6662 


240 


125 


-0.6573 


-0.6625 


-0.6685 


-0.6752 


-0.6829 


235 


130 


-0.6830 


-0.6852 


-0.6875 


.-0.6901 


-0.6927 


230 


[135 


-0.7030 


-0.7021 


-0.7009 


-0.6993 


-0.6970 


225 


140 


-0.7181 


-0.7142 


-0.7096 


-0.7040 


-0.6972 


220 


145 


-0.7292 


-0.7225 


-0.7137 


-0.7055 


-0.6944 


215 


150 


-0.7370 


-0.7279 


-0.7172 


-0.7047 


-0.6898 


210 


155 


-0.7423 


-0.7310 


-0.7178 


-0.7025 


-0.6843 


205 


160 


-0.7459 


-0.7326 


-0.7173 


-0.6997 


-0.6788 


200 


165 


-0.7480 


-0.7333 


-0.7163 


-0.6968 


-0.6738 


195 


170 


-0.7492 


-0.7334 


-0.7153 


-0.6944 


-0.6700 


190 


175 


-0.7498 


-0.7334 


-0.7146 


-0.6929 


-0.6675 


185 


180 


-0.7500 


-0.7333 


-0.7143 


-0.6923 


-0.6667 


180 



ENGINE DYNAMICS 



43 



The resultant pressure, P, curve C, acts along the axis of the 
cylinder. The force acting along the connecting rod, Fig. 32, 
is 

P E = P -f- cos b. 

The component acting tangentially to the 
crankpin circle is 

Pq = Pe sin (a + b) = P sec b sin (a + &)• 

Table 2 gives values of the tangential 
factor [sec b sin (a + &)]. 

The angles a and ?> are connected by the 
equation 

r . 



sin b 

Consequently, 

P Q = P sin a (l + 



sin a. 



r cos a 



\/£ 2 — r 2 sin 2 a ' 
The torque or turning moment applied 
to the crank at any crank angle a is 

T = P Q r 

Figure 33 shows the torque variation for a 
single cylinder of the Liberty 12 engine. Fig. 32. — Diagram 
Since the brake m.e.p. is 123 lb., the engine Rowing effect of obliq- 

uity of connecting rod. 

horse power per cylinder is 




123 X ~ X (I X 5 2 ) 
33,000 



X850 



= 36.8 



The mean torque at the propeller must lead to the same result : 
2tuT = 36.3 X 33,000, or, T = 112; that is, the mean torque 
per cylinder is 112 lb.-ft. The mean torque at the crankpin 
as determined from Fig. 33 is greater than this by the torque 
required to overcome the frictional resistance of one cylinder, 
or one-twelfth of the total frictional torque of the engine. If 
the mechanical efficiency of the engine is 85 per cent, the indicated 
mean effective pressure is 10 %5 X 123 lb. per square inch and 
the mean torque per cylinder is 10 %5 X 112 lb.-ft. The total 
horse power for the 12-cylinder engine is 12 X 36.3 = 436 and 
the mean total crankshaft torque is 12 X 112 = 1,345 lb.-ft. 



44 



THE AIRPLANE ENGINE 

Table 2. — Tangential Factors 

sin (a + b) 
cos b 



Crank 


I 


I 


I 


I 


'—3. 


Crank 


angle, 


- = 4 


- = 3.75 


- = 3.5 


- = 3.25 


angle, 


degrees 


r 


r 


r 


r 


r 


degrees 





0.0000 


0.0000 


0.0000 


0.0000 


0.0000 


360 


5 


0.1089 


0.1103 


0.1119 


0.1139 


0.1161 


355 


10 


0.2164 


0.2193 


0.2226 


0.2264 


0.2307 


350 


15 


0.3214 


0.3257 


0.3305 


0.3360 


0.3425 


345 


20 


0.4227 


0.4281 


0.4343 


0.4415 


0.4499 


340 


25 


0.5189 


0.5254 


0.5329 


0.5415 


0.5515 


335 


30 


0.6091 


0.6165 


0.6250 


0.6314 


0.6464 


330 


35 


0.6923 


0.7003 


0.7098 


0.7206 


0.7333 


325 


40 


0.7675 


0.7761 


0.7860 


0.7974 


0.8108 


320 


45 


0.8340 


0.8429 


0.8529 


0.8647 


0.8786 


315 


50 


0.8914 


0.9001 


0.9101 


0.9219 


0.9358 


310 


55 


0.9391 


0.9475 


0.9572 


0.9685 


0.9819 


305 


60 


0.9770 


0.9847 


0.9938 


1.0041 


1.0167 


300 


65 


1.0046 


1.0116 


1.0195 


1.0290 


1.0401 


295 


70 


1.0223 


1 . 0282 


1 . 0352 


1 . 0430 


1.0524 


290 


75 


1.0303 


1.0349 


1 . 0401 


1 . 0464 


1.0539 


285 


80 


1.0290 


1.0320 


1.0356 


1.0399 


1.0452 


280 


85 


1.0186 


1.0202 


1.0221 


1.0242 


1.0268 


275 


90 


1.0000 


1.0000 


1.0000 


1.0000 


1.0000 


270 


95 


0.9739 


0.9723 


0.9703 


0.9680 


0.9656 


265 


100 


0.9408 


0.9376 


0.9339 


0.9296 


0.9245 


260 


105 


0.9016 


0.8970 


0.8916 


0.8853 


0.8780 


255 


110 


0.8570 


0.8511 


0.8443 


0.8364 


0.8268 


250 


115 


0.8082 


0.8011 


0.7930 


0.7836 


0.7723 


245 


120 


0.7551 


0.7473 


0.7384 


0.7278 


0.7153 


240 


125 


0.6989 


0.6907 


0.6811 


0.6697 


0.6563 


235 


130 


0.6406 


0.6320 


0.6219 


0.6102 


0.5962 


230 


135 


0.5801 


0.5713 


0.5613 


0.5495 


0.5356 


225 


140 


0.5181 


0.5094 


0.4997 


0.4882 


0.4748 


220 


145 


0.4549 


0.4468 


0.4375 


0.4267 


0.4140 


215 


150 


0.3908 


0.3835 


0.3750 


0.3652 


0.3536 


210 


155 


0.3263 


0.3198 


0.3124 


0.3038 


0.2936 


205 


160 


0.2614 


0.2559 


0.2498 


0.2425 


0.2339 


200 


165 


0.1962 


0.1920 


0.1872 


0.1817 


0.1751 


195 


170 


0.1309 


0.1280 


0.1247 


0.1209 


0.1166 


190 


175 


0.0654 


0.0640 


0.0624 


0.0604 


0.0582 


185 


180 


0.0000 


0.0000 


0.0000 


0.0000 


0.0000 


180 



With the firing order used on the Liberty engine and with 
a Vee angle of 45 deg., the firing intervals between two cylinders 
on any one crank are 315 and 405 deg. of crank revolution. The 
turning moment on any one crank is obtained by superimposing 



ENGINE DYNAMICS 



45 



1000 



800 



600 



« 400 

c 
o 

cp200 



-200 



-400 



A 








1 
















\ 


























\ 




























\ 


























\ 


























\ 


























\ 


























\ 


L 










/^ 










We 


!7/7 


v° 


rqu 


? /I \ 




/ 


\ 












\ 










\ 






A 






\ 






11 






\ 


\ 




1 720 






















\ / 
























\ 































100 200 300 400 500 WO 100 
Crank Angle, Degrees 

Fig. 33. — Turning moment for a single cylinder of the Liberty- 12 engine 




100 200 300 400 500 600 700 

Crank Angle, Degrees 
Fig. 34. — Turning moment on each crank of the Liberty-12 engine. 



46 



THE AIRPLANE ENGINE 



two torque curves like those of Fig. 33 with a phase lag of 315 
deg. Adding the ordinates of two such curves, as in Fig. 34, 
gives the total torque on one crank. 

The six cranks of this engine are spaced at angular intervals 
of 120 deg. The total torque on the crankshaft is obtained by 
superimposing six curves like the resultant curve of Fig. 34, with 
angular intervals of 120 deg., and taking the algebraic sums of 
ordinates at the various crank positions. This process gives the 
curve of Fig. 35. The torques and torque ratios are as follows: 



One One 

cylinder crank 



Whole 

ENGINE 



Maximum crankshaft torque, pound-feet 1,030 1,240 1,670 

Ratio of maximum to mean torque 9.2 5 . 54 1 . 24 




Fig. 35. 



200 300 400 500 
Crank Angle, Degrees 

-Torque at propeller end of the crankshaft of the Liberty- 12 engine. 



The ratio of maximum to mean torque varies with the angle of 
the Vee. For 5- by 7-in. cylinders at 120 lb. mean effective 
pressure, the following ratios hold for torques on one crank: 



Vee angle 



2V 

T*. 



t5 Deg. 


60 Deg. 


75 Deg. 


90 Deg. 


5.2 


5.1 


5.6 


4.8 


-2.7 


-2.3 


-1.7 


-1.5 


7.9 


7.4 


7.4 


6.3 



Here 7\ = maximum torque -5- mean torque, 
T 2 = minimum torque -f- mean torque, 
T 3 = range of torque -s- mean torque. 

For the whole engine, ratios are as follows: 



8-cylinder. 



12-cylinder. 



Vee 


ANGLE 


Ti 


T 2 




[90 


1.40 


0.66 




75 


1.42 


0.18 


, 


60 


1.70 


-0.13 




, 45 


2.14 


-0.26 




f60 


1.13 


0.86 




1 45 


1.25 


0.89 



ENGINE DYNAMICS 



47 



In both tables negative signs indicate reversal of direction of tor- 
que. The minimum torque variation is seen to occur with equal 
firing intervals (eight-cylinder, 90-deg., and 12-cylinder, 60-deg. 
Vee engines). The great increase in this variation as the Vee 
angle of the eight-cylinder engine is diminished is very marked. 
A smooth curve of crankshaft turning moment, approximating 
as closely as possible to the mean torque line, is in every way 
desirable. This can best be obtained by the use of a plurality 
of cylinders with equal firing intervals. The greater the number 
of cylinders, the more uniform is the torque. The firing order 
of the cylinders is unimportant from this standpoint, as long as 
the firing interval is constant. The firing order is of the utmost 
importance, however, in relation to the balancing of forces and 
the stresses in the engine. Smoothness of running depends on 
the magnitude of the areas enclosed between the total torque 
curve and the mean torque line (Fig. 35). Areas above the line 
represent work done by the engine in excess of the resisting pro- 
peller torque and lead to acceleration; areas below the line 
represent a deficiency in engine work and a consequent slowing 
down. In ordinary engine practice a flywheel is used to absorb 
the excess and make up the deficiency without permitting 
excessive change in engine speed. In an airplane engine the 
propeller takes the place of a flywheel ; its large radius of gyration 
enables it to absorb a considerable amount of excess work with 
only a very small increase in speed of rotation. Moreover, the 
resisting torque varies as the square of the angular velocity and 
hence increases notably with small increases in rotative speed. 
The influence of the number of cylinders on the variation of 
crankshaft torque is clearly shown in the following table. 1 
The firing interval is constant in all cases. 



Table of Torque Variation 



Number of cylinders. . . 

Ratio of maximum in- 
stantaneous torque 
to mean torque 

Relative values of the 
maximum torque at 
propeller 



7.70 



1.0 



5.20 



1.35 



2.74 



1.07 



2.94 



1.64 



1 . 53 1 . 06 



1.17 



1.45 



0.91 1.32 



1.40 



1.22 



1.45;1.43 



10 



1.12 



1.46 



12 



1.13 



1.76 



16 



1.06 



2.20 



IS 



1.03 



2.41 



The relative values of the maximum torque are also given in 
the table, the value for a single cylinder being taken as unity. 
x From article by G. D. Angle, Aviation, Oct. 1, 1919. 



48 



THE AIRPLANE ENGINE 



It will be seen that the maximum torque at the propeller end of a 
6-cylinder engine is less than the maximum torque exerted by a 



1000 
< 800 

* m 

3 400 

i wo 
I o 

H -200 
TS -400 
£ -600 



■1000 



f=-s 


1 for:"::^-:::::z 


■hz±-zfttpi~-?^ 


•zzzzz^izfihz^t 





4000 ■; 
3000 sj 
2000 §_ 
1000 t 

.-# 

1000 8 

2000 ^ 
3000 | 
4000 (E 



180 



360 
Crank Ancjle, Degrees 



540 



720 



Fig. 36. — Side thrust against the cylinder walls of the Liberty-12 engine. 

single cylinder and consequently the crankshaft must be as 

strong at the free end as at the propeller end. 

Side Thrust. — Side thrust of the piston 
against the cylinder wall exists in con- 
sequence of the obliquity of the con- 
necting rod; it disappears at the two 
dead centers. Its magnitude, G (Fig. 32), 
is given by G = P tan b. Since r sin a = 
I sin 6, 

G _ r sin a 

p Vl 2 -r 2 sin 2 a 

For the Liberty engine with indicator 
diagram, as in Fig. 30, the side thrust is 
as shown in Fig. 36, the maximum value 
reaching nearly 1,000 lb. Change of sign 
indicates change of thrust from one side 
of the cylinder to the other. In sta- 
tionary cylinder engines side thrust is 
important only in relation to frictional 
wear. 

Offset Cylinders.— The side thrust 
during the expansion stroke, and the 
friction loss and cylinder wear, may be 
reduced by offsetting the cylinder: that is, 
by so locating it that its center line does not pass through the axis 
of the crankshaft. Minor effects of this arrangement are that 




Fig. 37. — Diagram show- 
ing the effects of the obli- 
quity of the connecting rod 
in an offset cylinder. 



ENGINE DYNAMICS 49 

the piston stroke is slightly greater than twice the crank throw 
and the mean speed of the piston is greater during the down 
stroke than the up stroke. This last point has the advantage 
of reducing the heat loss to the cylinder walls during the 
expansion period. 

In Fig. 37, if k = offset and x = distance from a point on the 
piston to a horizontal plane through the crankshaft axis (in 
a vertical engine),, then — 

x = r cos b -j- I cos a 

dx . , 7 . da 

do bd 

d 2 x 7 7- d 2 a , (da\ 2 

— — r cos o — I sm a ~tto ~~ * ( -ji 1 cos a 
db 2 db 2 \dbl 

. . (r sin b — k\ 
a =sln _i^ . j 

da r cos b 
db I sin a 

, 9 — I cos ar sin b + r cos bl sin a -^r 
a'a _ do 

db 2 I 2 cos 2 a 

d 2 x , , r 2 cos 2 b 

= — r cos o — t cos a 



2 Z 2 cos 2 a 

Z cos ar sin 6 + r cos bl sin a- 

+ Z sin a 



da 



Z 2 cos 2 a 

, r 2 cos 2 6 . , r 2 cos 2 6tan 2 a 

= — r cos o = h r sm o- tan a ; 

I cos a I cos a 

T 2 cos 2 b 

= — r cos b — i (1 + tan 2 a) + r sin 6-tan a. 

I cos a 

The acceleration of the reciprocating parts is equal to 

/2wn\ 2 d 2 x 2 d 2 x 

( 667 lb 2 = °- 011nt *. 

and the accelerating force is 

P.-f(0.011»^). 0.00034 Wg 

The side thrust is given by (z = P tan a. 

In using the above equations it should be noted that the angle 
a is positive when the connecting rod swings away from the 
crankshaft, as in the position shown in Fig. 37, and becomes 



50 



THE AIRPLANE ENGINE 



negative on the other side of the vertical. The angle a is found 
for any value of b from the equation 

I sin a = r sin b — k 




"0 20 40 60 80 100 120 140 160 180 Z00 220 240 260 280 300 320 340 360 
Degrees Rota+ion of Crank from Top Dead Cen+er 
Fig. 38. — Effects of different degrees of offset on the side thrust in a single 
cylinder of the dimensions of the Liberty-12 engine. 

The results of an analysis of the Liberty engine with offsets of 
0.5, 1.0 and 1.75 in. gives side thrusts as shown in Fig. 38. It 
will be seen that with an offset of half the crank throw (1.75 in.) 
the side thrusts during the exhaust and compression strokes are 

nearly as high as the maximum 
value reached in an engine with- 
out offset. The lowest maximum 
is obtained with an offset of 1 in. 
Rotary Engines. — The turning 
moment in a rotary engine results 
entirely from side thrust on the 
cylinder walls. This thrust is due 
not only to the obliquity of the 
connecting rod, as with stationary- 
cylinder engines, but also to thrust 
resulting from tangential acceleration of the reciprocating parts. 
The radial and tangential accelerations of these parts and the 
inertia forces resulting from them may be determined as follows : 

Take a single cylinder, as in Fig. 39, rotating about the shaft 
O, while the connecting rod rotates about the fixed crankpin P. 




Fig. 39. — Diagram of rotary engine. 



ENGINE DYNAMICS 51 

The piston pin Q will move along the axis OX as that axis 
rotates about Q. If the length OQ = x, the point Q will undergo 
radial acceleration a R along OQ and also tangential acceleration 
a T at right angles to OQ. The magnitudes of these accelerations 
are given by the general theorems: 

d 2 x /da\ 2 

OiR = 

and 



x -) 
dt 2 X \dt) 



n dx da . d 2 a 

aT = 2 Wdt +x W 

With rotary engines the angular velocity co of the cylinders is 
constant; or 

da 



and 

Consequently 

and 



dt="- 



dt 2 U ' 



d 2 x 



dx 
dt 



dx d 2 x 

To find values of ~r. and -37^ , the relations, r sin a = I sin b, 

and, x — r cos a + I cos b, are used. Combining these, 



, , L r 2 sin 2 a 
r cos a + Ul =— 



= r cos 



a + i (1 — — n™~) approximately. Then 



x I r . s 

- = cos aH ^- 7 sm 2 a 

r r 21 



But a = co£ 
therefore 



X ^ I * r • 2 4 

- = - + cos cot — ^ sm 2 co£ 
r r 21 

1 dx . r . 

— =r = — co sm cot — co ^ sm co£ cos cot 

r dt I 

1 d 2 x T 

- -37^ = — co 2 cos cotf + co 2 -, (sin 2 co£ — cos 2 wQ. 

r dt- I ' 



52 



THE AIRPLANE ENGINE 



Then, 



a R I d 2 x x 



— -,.,2 



r r dt 2 r 
I 



w 2 — 2co 2 cos cot + ~co 2 7 sin 2 cot — co V cos 2 cot 
r 2 1 I 



and 



<x T = 2co 



5 



2co 2 r sin cot (l + j cos con 



The values of a R , multiplied by the mass of the reciprocating 
parts, give the forces necessary to overcome inertia in the radial 
direction. Combining these forces with the total gas pressures 
(as in Fig. 31) gives the axial force P from which the side thrust 
G (Fig. 32) resulting from obliquity of the connecting rod is 
obtained, as on page 48. To find the total side thrust there 
must be added to this the thrust due to the tangential accelera- 
tion a T , which is the product of the acceleration by the mass 
of the reciprocating parts. The product of the total side thrust 
by the distance of the piston pin from the crankshaft is the 
turning moment. 

For several cylinders the turning moments are additive. As 
the cylinders are spaced at equal angular intervals, the total 
turning moment at any angular displacement a of one cylinder is 

T = T(a) + V„ L -°) + V°-™>) + • • • + V^ 1 -) 

The quantities in brackets 
are the angular displacements 
of the various cylinders: n is 
the total number of cylinders. 
Mayer 1 has calculated the 
turning moments for vari- 
ous arrangements of rotating- 
cylinder engines. For a seven- 
cylinder engine, 110 mm. diameter, 120 mm. stroke, length of 
connecting rod 213 mm., weight of reciprocating parts 1.3 kg., 
making 1,600 r.p.m., the radial and tangential accelerations a R 
and a T vary throughout the revolution as indicated in Fig. 40. 
It will be seen that the radial acceleration is always negative, 
*" Etude dynamique des moteurs d cylindres rotatifs." 




Fig. 40. — Variation of radial and tan- 
gential accelerations of the reciprocating 
parts of a rotary engine. 



ENGINE DYNAMICS 



53 



4000^ 



3000- 



eooo- 



1000- 



-1000- 



-eooo- 



-3000- 



I 

\ 



instead of changing sign as in the acceleration of the recipro- 
cating parts of stationary-cylinder engines. Combining the 
inertia force required to give the reciprocating parts this radial 
acceleration with the gas pressure force, the resultant force 
acting along the cylinder axis varies as shown in the solid black 
line in Fig. 41. It will be 
seen that this force is 
positive only for a very 
short period at the begin- 
ning of expansion. If the 
speed of the engine is 
increased the condition is 
soon reached when the re- 
sultant force on the piston 
pin is negative throughout 
the cycle. In that case 
the connecting rod would 
be under tension all the 
time (a very favorable con- 
dition, permitting consid- 
erable lightening of the 
rod) and the connecting 

rod might be replaced by FlG - 41.— Forces acting along the cylinder 
. r . axis of a rotary engine. 

a cham except that it 

is under compression before the engine has attained its full 

speed. 

The turning moment for a seven-cylinder Gnome engine of the 
dimensions given above is shown in Fig. 42. The maximum 
excess of power developed over the mean resistance (shaded 



\J80 



Degrees 
360 



540 







N 



w 



I 




\J 



<? 



I'.O 




360 540 720 

Degrees 

Fig. 42. — Turning moment of a 7-cylinder Gnome engine. 

area) is 60 ft.-lb., or about ;Ki20 of the total kinetic energy of the 
engine, and would be a much smaller fraction of the kinetic 
energy of engine and propeller combined. The earlier Gnome 
engines governed by cutting out the explosion on one or more of 
the cylinders according to the power requirements. The cutting 



54 THE AIRPLANE ENGINE 

out of a cylinder has a serious effect on the uniformity of turning 
moment and results in a maximum deficiency of work done 
during the cycle of 280 ft.-lb. as compared with the maximum 
excess (or deficiency) of 60 ft.-lb. when all cylinders are 
functioning. 

The resultant force on the crankpin of the above engine at 
full load varies from 2,640 to 800 lb. 

BALANCING 

The forces acting in an engine are of two kinds, in so far as 
their effect on stresses is concerned. The gas pressure in the 
cylinder, being exerted both on the cylinder head and on the 
piston, subjects the engine to equal and opposite forces which 
are inherently balanced, or, in other words, has no effect in 
displacing the engine as a whole. The inertia forces of the 
moving part in each cylinder are not inherently balanced. If 
the engine is to operate smoothly (without vibration) it must be 
so designed as to make these unbalanced forces counteract one 
another as far as possible. 

There are two kinds of moving parts to be considered: (a) 
the rotating parts, and (6) the reciprocating parts. 

Rotating Parts. — A body of mass m (weight W = mg) revolving 
with angular velocity co radians per second (n r.p.m.) in a circle 
of radius r feet has an acceleration of co 2 r which gives rise to a 
centrifugal force F = mco 2 r = 0.00034 Wn 2 r lb. acting radially. 
The rotating body is usually composed of the crank, crankpin, 
and the large end of the connecting rod. The resulting force F 
is an unbalanced force acting on the crankshaft. In multicrank 
engines there will always be a number of these forces, one acting 
at each crank. If there is more than one connecting rod attached 
to the crank (as in Vee and fixed radial engines) the revolving 
mass includes the large ends of all the connecting rods attached 
to the crank. In rotating cylinder engines the revolving mass 
includes the cylinders themselves but not the crank and 
crankpin. 

It is easy to balance the rotating parts. To accomplish this 
it is necessary (1) that the vector sum of these unbalanced forces 
should be zero and (2) that the sum of their couples about any 
(arbitrarily selected) plane should be zero. As the centrifugal 
forces on all the cranks are equal, the condition (1) is met, with 
any number of cylinders greater than one, by making the crank 



ENGINE DYNAMICS 55 

angle intervals equal. Condition (2) can be met, with any num- 
ber of cylinders greater than two, if the cranks are spaced at 
equal distances apart along the shaft. 

There is an unbalanced centrifugal pressure on the crankpin 
due to the inertia of the big ends of the connecting rods, and on 
the main bearings due to the inertia of all the rotating parts. 

Reciprocating Parts. — The inertia force of the reciprocating 
parts has already been seen to be 

P a = 0.00034Tfr? 2 r(cos a ± j-cos2a) 

If the connecting rod were infinitely long the expression would be 

Pa! = 0.00034 Wn 2 r-cos a. 

The second term in the brackets is' due to the obliquity of the 
connecting rod and increases in value as I becomes shorter. 
It is customary to separate the two terms in a discussion of 
balancing, the quantity P aI being called the primary inertia force, 
while 

P a n = ± 0.00034TFn 2 y-cos2a 

is called the secondary inertia force. It is found that the balance 
of the primary forces is much more easily achieved than that of 

T 

the secondary forces. Since j is usually about 34, the magnitude 

of the secondary forces is only about one-quarter that of the 
primary forces so that secondary balance is not as important as 
primary balance. 

The conditions for balance of primary and secondary inertia 
forces are exactly the same as those for centrifugal forces. There 
is, however, this difference, that these forces act always in the 
direction of the line of stroke. The practice of balancing the 
primary forces by the use of counterbalancing masses attached to 
the crankshaft is inadmissible (where avoidable) because of the 
necessity of keeping the total weight as low as possible and also 
because the same result can be obtained by multicylinder 
construction. 

The following general results of an analysis of various possible 
cylinder arrangements may be stated, it being assumed that the 
reciprocating masses are the same in all cylinders : 

With cylinders in a line and equally spaced at intervals of 
Lit.: 



56 THE AIRPLANE ENGINE 

For three-crank engines with cranks at 120 deg., the primary 
and secondary forces are balanced, but the primary and secondary 
couples are not. The maximum unbalanced primary couple is 
\/S'mo) 2 rL. 

For four-crank engines with cranks at 180 deg., the primary 
forces are balanced but the secondary forces are not. The 
secondary forces in a vertical engine have a vertical resultant 

which is equal to 4mco 2 y- at each quarter turn (0 deg., 90 deg., 180 

deg., etc.) and become zero at each eighth turn (45 deg., 135 
deg., etc.). 

For five-crank engines with cranks at 72 deg., the primary and 
secondary forces are balanced, but the couples are not. The 
maximum unbalanced primary couple is 2.5 moo 2 rL; the maximum 

unbalanced secondary couple is 5mco 2 y L. 

For six-crank engines with cranks at 120 deg., the primary and 
secondary forces and couples are all balanced. 

With opposed cylinder engines, with two cylinders and cranks 
at 180 deg., the primary and secondary forces are balanced; the 
primary couple is unbalanced with a maximum value of mu 2 rL. 

With Vee engines : 

For six cylinders, angle of Vee 120 deg., three cranks at 120 
deg., the primary and secondary forces are balanced; the couples 
are unbalanced; maximum unbalanced primary couple is 

V3. ■ T 
— — murrL. 

z 

For eight cylinders, angle of Vee 90 deg., four cranks at 180°, 

primary forces balanced, secondary forces unbalanced; primary 

and secondary couples balanced. Unbalanced secondary force 

acts always at right angles to the longitudinal plane of symmetry 

r 2 
and has a maximum value of 5.6 mco 2 y 

For twelve cylinders, angle of Vee 60 deg., six cranks at 120 deg., 
primary and secondary forces and couples all balanced. 

With fixed radial engine of k cylinders with all the connecting 
rods on one crank, the inertia forces have a constant resultant 

k 
of ~mco 2 r, approximately, along the crank. It can be balanced 

km 
completely by a counterbalance mass of — ~- at radius r, opposite 



ENGINE DYNAMICS 57 

the crankpin. If there are two banks of cylinders, each with k 
cylinders and with cranks at 180 deg., the primary forces will 
balance but there will remain an unbalanced primary couple of 

-•mco 2 rL. 

Rotating-cylinder engines have two sets of rotating masses: 

1. The cylinders, which are perfectly balanced, and 

2. The pistons and connecting rods, which have primary 
balance but not secondary balance. 

PERIODIC UNBALANCED FORCES 

The results of the existence of periodic unbalanced forces 
in an airplane engine may be two-fold: (1) to move the engine as 
a whole, and (2) to distort the engine. 

The engine in an airplane is supported on wooden members 
which are flexible and are in no way the equivalent of a rigid 
foundation. If the engine moves as a whole it will flex its supports 
instead of moving the airplane as a whole. The maximum 
possible duration of the periodic variation of any unbalanced 
force is one engine revolution or about J125 sec; ordinarily it will 
be not more than one-quarter of a revolution or Jioo sec. This 
time is too short to permit the unbalanced force to produce 
appreciable deviation of the airplane as a whole although it 
may set up vibration in it. 

The crankcase of the airplane engine is always very thin 
and consequently flexible. Much of the unbalanced force may be 
taken up in producing distortion of the crankcase. This is 
very undesirable since it necessarily results in varying align- 
ments of the crankshaft bearings and consequent increase in 
shaft friction. 

Both the engine supports and the crankcase have their own 
natural periods of vibration. If the frequency of the distur- 
bances due to unbalanced forces in the engine is the same as the 
natural frequency of the vibration of either of these members, the 
amplitude of vibration will be greatly increased beyond the very 
small amount due to a single application of the unbalanced force. 
It is most important that the critical speed at which the vibration 
of the engine on its supports becomes a maximum should be 
avoided, as well as simple multiples of those speeds. The natural 
periods of vibration of such complex structures as engines, on 
their supports in an airplane, can not be calculated. If excessive 



58 



THE AIRPLANE ENGINE 



vibration is found at or near the designed speed of the engine 
the structures must be altered to change the natural period. 
Various devices have been used for neutralizing the unbalanced 
secondary forces in four- and eight-cylinder engines. A good 
example of these is the Lanchester balancer (Fig. 43), which has 
had considerable application to automobile engines. This 
consists of two exactly similar unbalanced cylinders located under 
the center main bearing and driven by a gear on the crankshaft. 
These cylinders revolve in opposite directions and their unbalanced 
weights are located so that their common center of gravity travels 
up and down in a vertical plane and so balances the displacement 
of the common center of gravity of the pistons, which falls in a 
plane at the middle of the piston travel when the pistons are on 




Mid-stroke 
Displacement 
- K . due to Angu- 
' larityofffods 



Fig. 43. — Lanchester balancer. 



their dead centers, but falls below it when the cranks have revolved 
90 deg. past that position. 

It should be clearly recognized that complete balance of 
primary and secondary forces and couples does not ensure absence 
of vibration. Such complete balance means that the engine as a 
whole has no tendency to move, but there are always internal 
stresses, particularly those imposed by the opposing couples on 
the engine structure. If the periodicity of the application of these 
stresses coincides with the natural period of the structure (or 
some fraction of it) severe vibrations may be set up. In any case 
heavy bearing loads are likely to be imposed, especially in the 
center main bearing, through the action of opposing couples. 

For torsional oscillations of the crankshaft see p. 146. 

The following table gives values of inertia and centrifugal 
forces, resulting bearing pressures and other calculated quantities 



ENGINE DYNAMICS 



59 



for the six-cylinder 200-h.p. Austro-Daimler engine, the six- 
cylinder 270-h.p. Basse-Selve engine and the 12-cylinder 400-h.p. 
Liberty engine: 



Inertia Forces, Bearing Loads, Etc. 



Austro- 
Daimler 
engine 



BassS- 
Selve 
engine 



Liberty- 

12 
engine 



Weight of piston complete with rings and piston pin, lb. 

Weight per sq. in. of piston area, lb 

Weight of connecting rod complete, lb 

Weight of reciprocating part of connecting rod, lb 

Total reciprocating weight per cylinder, lb 

Weight per sq. in. of piston area, lb 

Length of connecting rod (centers), in 

Ratio connecting rod length to crank throw 

Inertia, lb. per sq. in. piston area (top center) 

Inertia, lb. per sq. in. piston area (bottom center) 

Inertia, lb. per sq. in. piston area (mean) 

Weight of rotating mass of connecting rod, lb 

Total centrifugal pressure, lb 

Centrifugal pressure, lb. per sq. in. piston area 

Mean average loading on crankpin bearing, total from 

all sources, lb. per sq. in. piston area 

Diameter of crankpin, in 

Rubbing velocity, ft. per sec 

Effective projected area of big-end bearing, sq. in 

Ratio piston area to projected area of big-end bearing . 
Mean average loading on big-end bearing, lb. per sq. in. 



4.18 

0.188 

4.84 

1.66 
5.84 
0.263 

12.40 
3.6:1 

63.8 

36.2 

25.0 
3.18 

610.0 
27.5 

91.0 
2.20 
13.42 
5.02 
4.42:1 
402.0 



6.187 
0.211 
9.00 

2.25 
8.437 
0.288 
14.17 
3.6:1 
80.7 
45.7 
31.6 
6.75 

1,480.0 
50.7 



115. 

2. 
16. 

5. 

3. 
405. 



7 

75 

8 

39 

5:1 





3.838 
0.1955 
2.9 and 
6.35 

225 

063 

258 



44:1 



2 



1 
5 


12 

3 

117 

65 



1 . 675 and 
5.125 
1,469.0 
74.8 

175.0 
2.375 

17.7 
5.34 
3.68:1 
642.0 



CHAPTER IV 

ENGINE DIMENSIONS AND ARRANGEMENTS 

Certain special requirements control the selection of the 
dimensions and arrangements of airplane engines. These are: 

1. Minimum total weight of the engine and its accessories per brake 
horse power developed, or maximum power output per pound of weight. 

2. Maximum fuel economy. 

3. Compactness. 

4. Freedom from unbalanced forces and from vibration. 

5. Reliability. 

To obtain minimum weight per horse power, it is necessary that 
the engine should have minimum weight per cubic foot of piston 
displacement per revolution, and that it should be operated with 
maximum power per cubic foot of cylinder volume. The latter 
demands a combination of the maximum obtainable mean 
effective pressure with high speed of revolution. The mean 
effective pressure is constant at moderate engine speeds but falls 
off at high speeds in consequence of falling volumetric efficiency. 
Beyond a certain limiting speed the mean effective pressure will 
fall off more rapidly than the increase in engine revolutions per 
minute (see p. 37) and the engine power will decrease. This 
limiting speed should be made as high as possible by making 
the valve openings large and the inlet and exhaust manifolds short 
and of ample cross-section. 

All airplane engines must be multicylindered in order to give 
the necessary uniformity of turning moment and freedom from 
unbalanced forces. The weight of the engine per cubic foot of 
piston displacement per revolution will depend on the unit size 
of cylinder selected. Comparing two unit cylinders of exactly 
the same form but of different sizes, it will be found that the 
thickness of the cylinder walls will not have to be increased as 
rapidly as the cylinder diameter because the wall (for structural 
reasons) is always made thicker than the stresses demand, by an 
amount which does not vary much with the diameter. Conse- 
quently the weight of the cylinder per cubic foot of piston dis- 

60 



ENGINE DIMENSIONS AND ARRANGEMENTS 61 

placement diminishes with increased size, and the same is true 
of most of the other engine parts. 

On the other hand, the weight of the engine per cylinder 
diminishes with increase in the number of cylinders in line. The 
engine consists of a number of exactly similar units (cylinder, 
running parts, section of crankcase, etc.) and certain approxi- 
mately constant weights such as ends of crankcase, pumps, 
magnetos, propeller hub and so forth. The addition of more 
cylinders will diminish the weight of the engine per cylinder. 

A further diminution in weight of the engine per cylinder 
can be obtained by using the Vee or W arrangement of cylinders. 
A substantial saving in the weight of crankshaft and crankcase, 
per cylinder, results from these arrangements. Still greater 
saving results from the adoption of the radial arrangement. 

Considerations of torque and balance (see p. 56) indicate 
that the number of cylinders should not be less than six for 
an engine of moderate power (100 to 150 h.p.). For higher 
powers the choice is between more cylinders and larger cylinders. 
Engines have been built with as many as 24 cylinders, but it does 
not seem likely that this number will be much used or exceeded. 

Cylinder size can be increased by increasing either diameter or 
stroke or both. The diameter is at present limited to from 6 to 
7 in. by the difficulty of keeping the piston cool. The heat 
given to the center of the piston has to travel radially to the 
walls, and it is necessary to increase the thickness of the piston as 
the diameter increases in order to give sufficient section of metal 
to carry the heat away. This results in a heavy piston and 
excessive inertia forces in the reciprocating parts. The engine 
stroke is also hmited at present to 8 in.; increase in stroke beyond 
that limit increases the over-all height of the engine to dimensions 
which are difficult to accommodate without increasing the size 
of the fuselage. Furthermore, increase in stroke increases the 
weight of the engine more than does a corresponding increase in 
diameter. The small ratio of stroke to diameter which character- 
izes airplane engines is not as objectionable as it would be in 
other engines; it increases the ratio of water-jacketed surface to 
cylinder volume, but the percentage of heat lost to the jacket is 
nevertheless smaller than in other engines in consequence of the 
high engine speed and high mean effective pressure. 

The weight of the engine per cubic foot of piston displacement 
per revolution is also a function of the ratio of connecting rod 



62 THE AIRPLANE ENGINE 

length to stroke. The smaller this ratio the less is the over-all 
height and the weight of the engine. The objection to a small 
ratio is the increase in magnitude of the secondary inertia forces 
which result from the obliquity of the connecting rod. These 
secondary forces can be perfectly balanced with certain arrange- 
ments of cylinders (see p. 56) and the objection eliminated. 
The ratio usually ranges from 1.5 to 1.7. 

High fuel economy is important primarily in its effect on the 
weight to be carried by the plane. For a five-hour flight, an engine 
weighing 2.5 lb. per horse power and using 0.5 lb. of fuel per horse- 
power hour will have the same total weight of engine and fuel 
as an engine weighing 2 lb. per horse power and using 0.6 lb. of 
fuel per horse-power hour. The heavier and more efficient 
engine would ordinarily be the better of the two in respect to 
reliability and durability. To obtain high fuel economy the 
most important factor is the ratio of compression which should 
be as great as can be used without detonation or preignition. 

Compactness is important both in respect to frontal area and 
over-all length. The frontal area of large engines should, if 
possible, be of such form and dimensions as not to require any in- 
crease in the cross-section of the containing fuselage. Short over- 
all length is distinctly advantageous and permits the fuel tanks 
to be located close to the center of gravity of the plane. 

Freedom from vibration is necessary because the mounting 
of the engine is on elastic supports — usually wooden longerons — 
which leave the crankcase free to distort under internal forces. 
Such arrangements of cylinders as will eliminate or minimize 
unbalanced forces are desirable but they cannot be relied upon 
to prevent vibration. Even with the most perfect balancing, 
torsional vibration of the crankshaft may produce excessive vibra- 
tions at certain engine speeds; such speeds must be avoided. 

Engine Arrangements. — Airplane engines have been built in a 
great variety of arrangements of which a number have survived. 
These may be classified as (1) radial, (2) vertical, (3) Vee, (4) 
W, (5) X. Radial engines (both fixed and rotary) are discussed 
in Chapter VIII. They have minimum weight per horse power 
and shortest over-all length, but they have maximum frontal 
area and until very lately have shown economy inferior to that of 
the other types. They are generally air cooled. 

The vertical engine (Figs. 8 and 9) has one row of cylinders. 
The Vee engine (Figs. 46 and 47) has two linear rows of cylinders; 



ENGINE DIMENSIONS AND ARRANGEMENTS 



63 



the " angle of the Vee" is the acute angle between the axial planes 
of the two rows. The W or \|/ engine (Figs. 72 and 73) has 
three rows of cylinders of which the central one is vertical and the 
other two form equal angles with the vertical. The X engine has 
four rows of cylinders arranged symmetrically about the vertical 
and horizontal planes but not necessarily with equal angles 
between the planes of the cylinder axes. In the radial engine 
(Fig. 136) three or more cylinders, constituting a group, have their 
axes intersecting at a point on a common shaft. The radial or 
fixed engine should not be confused with the rotary engine, in 
which the cylinders revolve about a stationary crankshaft. 
A radial engine may consist of more than one group or bank of 
cylinders, each group having its own common plane of cylinder 
axes, perpendicular to the axis of the shaft. 

Each cylinder of a four-cycle engine requires two revolutions 
or 720 deg. of rotation of the crankshaft to complete its cycle. 




fi 






n| 


j3l 


Ml 


Ml 




fH!Hte"f 




Fig. 44. 
Four-cylinder engine 



Fig. 45. 

Six-cylinder engine 



The explosions in a cylinder occur every other revolution. If 
explosions are to occur in n cylinders at equal intervals, the 
interval expressed in degrees of crankshaft rotation must be 
720 -f- n. Explosion always occurs when the piston is closest to 
the cylinder head. With a four-cylinder engine, the interval 
between explosions is 180 deg.; the crankpins lie all in one plane 
passing through the shaft axis, one possible arrangement being 
shown in Fig. 44. A six-cylinder vertical engine has the cranks 
720 -?- 6 = 120 deg. apart: the crankpins lie in three planes 
intersecting at the shaft axis. One arrangement is shown by 
Fig. 45. Constancy of interval between impulses may be 
obtained with crank dispositions other than those shown in 
Figs. 44 and 45. For example, in Fig. 44, cranks 2 and 3 may 
be either in line or opposed. If the former, cranks 1 and 4 will 
be in line and 180 deg. away from 2 and 3. If the latter, 1 and 4 
will be opposed, and either may be in line with 2. If equal inter- 



64 THE AIRPLANE ENGINE 

vals of time between explosions were the sole requisite, the 
number of possible crank arrangements with a 6-cylinder engine 
would be large. The actual crank arrangements are determined 
mainly by considerations of engine balance. 

In a Vee engine each pair of cylinders (in one plane) acts on a 

common crankpin. If there are n cylinders in ~ pairs, the crank 

interval is 720 4- ^. The interval between explosions of the 

cylinders is 720 -f- n. If 6 is the angle of the Vee, any one crank 
moves through the angle 6 in the interval necessary for the two 
pistons actuating it to reach respectively their highest positions. 
As the explosions occur with the pistons in their highest positions, 
the explosion in a leading cylinder will precede that of its fol- 
lowing cylinder, by an angle of 6, or 360 + 6, deg. according to 
whether both cylinders explode during the same revolution of the 
crank or the explosions occur in succeeding revolutions. The 
latter is always the case, because, if the second explosion occurred 
after the crank angle 6, the explosion pressure would be trans- 
mitted to a crankpin which was already subjected to the pressure 
of the gas expanding in the other cylinder of the pair and the 
crankpin would be unduly loaded. For equal explosion intervals 
the angle 6 is equal to 720 4- n. This leads to a 90-deg. angle 
for 8-cylinder and a 60-deg. angle for 12-cylinder Vee engines. 
These angles give the maximum uniformity of turning movement; 
other angles may be employed for special reasons but always at 
some sacrifice of uniformity of turning movement. 

The choice as between vertical, Vee and W arrangement is 
largely determined by the number of cylinders. It has been 
found that much trouble is experienced with a crankshaft having 
more than six cranks and the over-all length of the engine becomes 
excessive. Eight-crank engines have been built but they have 
not survived. The perfect balancing of the six-crank engine 
has made it a favorite. With six cylinders the vertical arrange- 
ment is almost universally adopted; it has the advantage of 
having minimum frontal area and consequently of being most 
easily accommodated in the fuselage. With eight cylinders the 
90-deg. Vee engine with four cranks is generally accepted as the 
best arrangement although the wide Vee angle results in con- 
siderable over-all engine width. With 12 cylinders the usual 
arrangement is a 60-deg. Vee and six cranks. The 45-deg. Vee 



ENGINE DIMENSIONS AND ARRANGEMENTS 65 

adopted in the Liberty engine results in decreased width of 
engine but increased height; it also results in a less uniform turn- 
ing moment on the crankshaft. 

The Warrangement for 12 cylinders is with three rows of cylinders 
and four cranks; this shortens the over-all length and decreases 
the weight but greatly increases the engine width, especially 
if a 60-deg. angle (which gives most uniform torque) is used 
between the rows. With 18 cylinders the W arrangement with 
three rows of cylinders and six cranks is used; the angle 
between the rows for most uniform torque should be 40 deg., 
which diminishes the over-all width as compared with the 12- 
cylinder W engine. 

Vertical, Vee and W engines are nearly always water- 
cooled in airplane practice. Air cooling has been successful only 
with low compression ratios. 

Engine Dimensions. — Table 3 gives the general dimensions 
and arrangement of the principal American and foreign airplane 
engines. The horse powers given are generally the maker's 
rating, but they are naturally variable with the fuel, the car- 
buretor, the manifolding, the ratio of compression, and the engine 
speed. The ratio of compression is readily varied by changing 
the dimensions of the piston and for certain engines different 
pistons are supplied, according to whether the engine is to be 
flown at low altitude (as in seaplanes) or at higher altitudes. 
The dry weight includes carburetors, magnetos, and propeller 
hub. The weight per horse power naturally varies with the 
horse power and is given for the rated horse power. It is the 
product of two factors, the weight per cubic inch of piston 
displacement and the piston displacement per horse power. 
The piston displacement per horse power is the aggregate dis- 
placement volume of the cylinders per stroke divided by the 
horse power developed. It is an excellent measure of the degree 
to which the piston displacement is utilized. The aggregate 
displacement volume is equal to I X a X n cu. in. where I is the 
stroke in inches, a the piston area in square inches, and n is the 

, ' „ , ^ , pX iXaXn XN 

number ol cylinders. Ine horse power = o v 19 y 33 qqq > 

where p is the mean effective pressure, and N is the revolutions 
per minute. The piston displacement per horse power = 33,000 
X 24 4- p X N; that is, it depends only on the mean effective 
pressure and the revolutions per minute. 



66 



THE AIRPLANE ENGINE 



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ENGINE DIMENSIONS AND ARRANGEMENTS 



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68 



THE AIRPLANE ENGINE 






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ENGINE DIMENSIONS AND ARRANGEMENTS 



69 



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70 



THE AIRPLANE ENGINE 



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71 






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72 



THE AIRPLANE ENGINE 

Table 4. — Detailed Dimensions of 



Engine 



Approximate horse power 



Liberty 



200 



400 



Packard 



180 



270 



Bore, inches 

Stroke, inches 

Number of cylinders 

Stroke-bore ratio 

Piston displacement per cylinder, cu. in 

Total piston displacement, cu. in 

Compression volume, cu. in 

Total volume of cylinder, cu. in 

Compression ratio 

Per cent compression 

Camshaft bearings, number on shaft . . . 
Diameter, inches 

Inlet valves: 

Number per cylinder 

Port diameter, inches 

Lift, inches 

Angle of seat 

Total area of opening, square inches . 

Tappet clearance, inches 

Exhaust valves: 

Number per cylinder 

Port diameter, inches 

Lift, inches 

Angle of seat 

Total area of opening, square inches . 

Tappet clearance, inches 

Valve springs: 

Number per valve 

Tension inlet (both springs): 

Valve open, pounds 

Valve closed, pounds 

Tension exhaust (both springs): 

Valve open, pounds 

Valve closed, pounds 

Internal spring tension: 

Valve closed, pounds 

Valve open (inlet), pounds 

Valve open (exhaust), pounds 

Valve timing: 
Inlet: 

Opens, past top center 

Closes, past bottom center 

Exhaust: 

Opens, before bottom center 

Closes, past top center 

Piston: 

Type 

Area of head, square inches 

Material 

Distance from center of pin to top of 

piston, inches 

Length over all, inches 

Length of bearing in cylinder 

Clearance in cylinder: 

Top, inches 

Bottom, inches 



5.00 
7.00 

6 
1.40 
137.4 
824.5 
31.1 
168.5 
5.42 
18.46 
7 
J 6-0.998 
\ 1-1.1235 

1 
2.50 
He 
30° 
3.202 
0.013-0.016 

1 
2.50 
0.375 
30° 

2.748 
0.019-^-0.021 



62.2 
50.0 

85.7 
71.5 

26.5 
32.9 
31.8 



10° 
45° 

48° 
8° 

Ribless 

19.63 
Aluminum 

3.469 
5.469 
3.75 

0.0395 
0.020 



5.00 
7.00 
12 



1 

137 

1650 

31 

168 

5 



40 

4 



1 

5 

42 
18.46 

7 
6-0.998 
1-1 . 1235 

1 
2.50 
He 
30° 
3.202 
0.013-0.016 

1 
2.50 
0.375 
30° 
2.748 
0.019-0.021 



62.2 
50.0 

85.7 
71.5 

26.5 
32.9 
31.8 



10° 
45° 



48° 



Ribless 

19.63 
Aluminum 

3.469 
5.469 
3.75 

0.0395 
0.020 



4.75 
5.25 

8 
1.105 
93.0 
744.0 
23.17 
116.2 
5.02 
24.9 

5 
4-0.998 
1-1 . 124 

1 
2.00 
He 
30° 
2.61 
0.013-0.016 

1 
2.00 
0.375 
30° 
2.20 
0.019-0.021 



10° 
45° 

48° 
8° 

Ribless 

17.72 
Al. alloy 

2.25 
4.25 
3.094 

0.037 
0.017 



4.75 
5.25 
12 

1.105 

93.0 

1116.4 

20.4 

113.4 

5.56 

18.0 

7 
6-0.998 
1-1.124 

1 
2.00 
He 
30° 
2.61 
0.013-0.016 

1 
2.00 
0.375 
30° 
2.20 
0.019-0.021 



63.5 
47.5 

81.5 
61.5 

24.0 
35.0 
35.0 



10° 
45° 

48° 
8° 

Ribless 

17.72 
Al. alloy 

2.67 

4.665 

2.99 

0.034 
0.017 



ENGINE DIMENSIONS AND ARRANGEMENTS 
American and German Engines 



73 



Hispano-Suiza 


Benz 


Maybach 


Austro- 
Daimler 


Basse- 
Selve 


Mercedes 


180 


300 


200 


300 


200 


270 


180 


4.72 


5.511 


5.512 


6.50 


5.31 


6.10 


5.51 


5.11 


5.905 


7.480 


7.09 


6.89 


7.87 


6.30 


8 


8 


6 


6 


6 


6 


6 


1.08 


1.07 


1.357 


1.09 


1.296 


1.29 


1.14 


89.9 


140.8 


178.4 


235.3 


152.8 


230.2 


150.3 


719.0 


1126.0 


1070.4 


1410.0 


916.8 


1381.0 


901.7 


20.7 


32.6 


37.2 


47.6 


38.0 


69.0 


41.3 


110.6 


173.4 


215.6 


282.9 


190.8 


299.2 


191.6 


5.33 


5.32 


5.8 


5.95 


5.02 


4.34 


4.64 


18.76 


18.8 


17.25 


16.8 


19.9 


23.0 


21.5 


3 


3 


4 


5 


4 






1.342 


1.344 


0.984 










1 


1 


2 


2 


2 


2 


1 


1.968 


2.205 


1.693 


1.89 


1.73 


2.20 


2.677 


0.393 


0.511 


0.433 


0.372 


0.390 


0.390 


0.453 


45° 


45° 


30° 


30° 


45° 


45° 




1.894 


2.79 


2.214 


2.33 


2.12 


2.72 


3.81 


0.078 


0.030 


0.016 


0.012 


0.01 




0.017 


1 


1 


2 


2 


2 


2 


7 


1.968 


2.205 


1.693 


1.89 


1.73 


2.20 


2.677 


0.393 


0.511 


0.433 


0.368 


0.40 


0.390 


0.453 


45° 


45° 


30° 


30° 


45° 


45° 




1.894 


2.79 


2.214 


2.33 




2.72 


3.81 


0.078 


0.030 


0.016 


0.016 


0.012 




0.014 


2 


2 


1 


1 


1 


1 


1 


75.0 


80.0 


43.0 


133.0 








44.5 


42.0 


25.5 


101.8 








75.0 


80.0 


44.0 


133.0 








44.5 


42.0 


24.0 


101.0 








19.0 














30.0 














30.0 














10° 


-10° 


5° 


-8° 


-10° 




0° 


50° 


62° 


45° 


35° 


30° 




40° 


45° 


62° 


55° 


33° 


45° 




40° 


10° 


29° 


18° 


7° 


7° 




10° 


Ribless 


Ribless 


Ribbed 


Ribless flat- 
top 


Ribbed 


Convex 
crown 


Steel concave 
crown 


17.53 


23.82 


23.82 


33.15 


22.2 


29.2 


23.84 


Aluminum 


Aluminum 


Al. alloy 


Cast iron 


Aluminum 


Aluminum 


Cast iron 


1.765 


3.00 


2.480 


3.19 


2.09 






4.375 


5.12 


4.843 
3.071 


5.944 


4.35 


5.11 




0.040 


0.040 


0.025 


0.029 


0.039 


0.020 




0.0165 


0.020 


0.008 


0.009 


0.016 


J 0.01 





74 



THE AIRPLANE ENGINE 



Table 4.— 



Engine 



Approximate horse power 



Liberty 



Packard 



200 



400 



180 



270 



Rings: 

Number per piston: 

Top 

Bottom 

Tension, pounds 

Width, inches 

Width of gap (ring in cylinder) 

inches 

Pin: 

Length, inches 

Diameter, inches 

Connecting rods, plain: 

Length c. to c 

Number of bolts 

Diameter of bolts, inches 

Thread 

Connecting rods, forked: 

Length c. to c 

Number of bolts 

Diameter, inches 

Thread 

Rod-stroke ratio , 

Wristpin bearing: 

Length, inches 

Diameter, inches 

Material 

Big end bearing, forked: 

Length, inches 

Diameter, inches 

Material 

Big end bearing, plain: 

Length, inches 

Diameter, inches 

Material 

Crankshaft bearings: 

Number 

Length, inches 

Diameter, inches 

Carburetor: 

Name 

Model 

Type 

Diameter of flange, inches 

Choke, inches 

Jets, diameter: 

Main, inches 

Compensating, inches 

Idling well or pilot jet, inches 



3 


ii-15 

0.248 

0.031 

4.234 
1.25 

12.0 

4 

0.311 
%6-24 



1.714 

2.00 

1.25 

Bronze 

babbitt 

lined 



2.25 

2.375 

Bronze 

shell babbitt 

lined 

7 
6-1 . 625 
1-4.25 



2.625 

Zenith 

55 AS 

Single nozzle 



2.165 
1.575 

1.80 mm. 
1 . 55 mm. 
1.00 mm. 



3 

11-15 
0.248 

0.031 

4.234 
1.25 

12.0 

2 
0.4365 
He-20 



12.0 
4 
0.3115 

%6-24 
1.714 

2.00 

1.25 

Bronze 



2.25 

2.375 

Bronze 

babbitt 

lined 



7 
6-1 . 625 
1-4.25 



2.625 

Zenith 

r. S., No. 52 

Double 

annular 

nozzles 

2.047 

1.417 

1 .65 mm. 
1.70 mm. 
1 . 00 mm. 



3 


9.0 
0.217 

0.031 

4.141 
1.25 

9.0 
2[ 
0.437 
Ke-20 



9.0 
4 
0.312 

2*6-24 
1.714 

2.00 

1.25 

Bronze 



2.492 
2.125 

Bronze 

babbitt 

lined 

1.125 
2.631 

Steel 



5 
1-1 . 625 
2-1.750 
1-2.625 
1-4.438 
2.375 

Packard 



3 


9.0 
0.217 

0.031 

4.138 
1.25 

9.0 

2 
0.437 
Ke-20 



9.0 

4 
0.312 

2*6-24 

1.714 

2.00 

1.25 

Bronze 



2.492 
2.125 

Bronze 

babbitt 

lined 

1.125 
2.631 

Steel 



7 
1-1 . 625 
4-1.750 
1-3.000 
1-4.438 
2.375 

Packard 



Double 


Double 


Venturi 


Venturi 


2.0 


2.0 


1.34 


1.42 



0.0730 
0.0700 
0.0785 



0.1285 
0.0995 
0.1040 



ENGINE DIMENSIONS AND ARRANGEMENTS 



75 



Continued 



Hispano-Suiza 



180 



300 



Benz 



200 



Maybach 



300 



Austro- 
Daimler 



200 



Selve 



270 



Mercedes 



180 



3 

1 
4.0 
0.0984 

0.011 

4.125 
1.181 

8.858 

2 

8.0 mm. 

Pitch, 

1 . 52 mm. 

8.858 

4 

12 mm. 

Pitch, 

1.25 mm. 

1.73 

2.203 

1.183 

Bronze 



2.508 

1.97 

Bronze 

babbitt 

lined 

1.25 
2.50 

Steel on 
bronze 



5 

1-3.86 

3-1 . 578 

1-Steel ball 

2.282 



Stromberg 
NAD 4 
Double 
Venturi 



2.18 
1.50 



0.0935 
0.1286 



1 
14.0 
0.25 

0.018 

4.781 
1.375 

10.375 
2 
0.500 



10.375 

4 
0.375 



1.76 

2.250 
1.375 
Bronze 



2.50 
2.125 
Bronze 
babbitt 
lined 

1.125 
2.75 
Steel on 
bronze 



5 

1-4.34 

3-2.00 

1-Steel ball 

2.50 



Stromberg 
NAD 6 
Double 
Venturi 

2.375 
1.812 



3 

1 
7.0 
0.118 

0.072 

4.764 
1.299 

12.992 
4 
0.394 
Pitch, 
0.9 mm. 



1.736 

2.795 
1.299 
Bronze 



0.116 



2.441 
2.362 
Bronze 
babbitt lined 



7 

1-1.772 

5-1 . 693 

1-2.402 

(1) -1.890 

(2-7)-2.441 

Benz 

BZ 3A-137 

Barrel 

throttle 



1.89 
1.654 



0.039 
0.020 



3 and one 

scraper 





0.255 

0.055 

6.26 
1.496 

12.205 
4 
0.551 



1.72 

3.66 
1.496 
Cast iron 
bush 



2.893 
2.598 
Bronze, 
white metal 



7 
2.638 



2.598 



Maybach 



Variable 



3 


0.275 

0.019 

5.10 
1.100 

12.40 

4 

0.39 

Pitch, 

1 . 5 mm. 



1.80 

2.64 
1.10 
Phosphor- 
bronze 



2.63 

2.20 

Bronze, 

white metal 



7 

1-2.20 

1-1.71 

5-1.97 

2.28 



Austro- 
Daimler 



Dual 



0.945 



0.230 



1.37 

14.17 

4 

0.47 

Pitch, 

1.75 mm. 



1.80 

3.14 
1.37 
Phosphor- 
bronze 



3.03 
2.75 
Bronze, 
white metal 



7 

1-2.36 

5-2.48 

1-3.74 

2.75 



Selve 



1.96 

0.0102 

0.0046 



Mercedes 

Twin jet dual 

0.945 
1.473 mm. 
0.558 mm. 



76 



THE AIRPLANE ENGINE 



Table 4. — 



Engine 


Liberty 


Packard 


Approximate horse power 


200 


400 


180 


270 


Ignition: 










Type 


Battery 


Battery 


Battery 


Battery 


Manufactured by 


Delco 


Delco 


Delco 


Delco 


Maximum spark advance, before top 










center 


30° 


30° 


45° 


45° 


Maximum retard, after top center. . . 


10° 


10° 


10° 


10° 


Speed ignition drives (CS = crank- 










shaft speed): 










Generator or magneto 


1.0 xcs 


1.5 XCS 


2.0XCS 


2.0XCS 


Distributors 


0.5XCS 


0.5XCS 


0.5XCS 


0.5XCS 


Firing order 


1-5-3-6-2-4 


1L-6R— 5L- 
2R-3L-4R- 


1L-4R-3L- 
2R-4L-1R- 


1L-6R-5L- 




2R-3L-4R- 






6L-1R-2L- 


2L-3R 


6L-1R-2L- 






5R-4L-3R 




5R-4L-3R 


Rotation, direction of (facing pro- 


Counter 


Counter 


Counter 


Counter 


peller) 


clock 


clock 


clock 


clock 


Spark plugs: 










Number per cylinder 


2 


2 


2 


2 


Location 


Cyl. head 


Cyl. head 


Side of head 


Side of head 


Size, millimeters 


18.0 


18.0 


18.0 


18.0 


Pitch, millimeters 


1.5 


1.5 


1.5 


1.5 


Gap, inches 


0.015-0.018 
1 . 5 X CS 


0.015-0.018 
1 . 5 X CS 


0.015 
1.5 X CS 


0.015 


Oil pump speed 


1.5 X CS 


Water pump speed 


1.5 X CS 


1.5 X CS 


1.5 X CS 


1.5 X CS 


Reciprocating and centrifugal 










weights : 










Piston, complete with rings and 










pin, pounds 


4.7 


4.9 


3.9 


9.2 


Upper end connecting rod, pounds. 


1.1 


1.3 


1.1 


1.1 


Lower end forked connecting rod, 










pounds 




4.4 


3.3 


3.3 


Lower end plain connecting rod, 










pounds 


3.7 


1.9 


1.4 


1.4 


Total weight of connecting rods 










complete, pounds: 










Forked 


0.0 


5.7 


4.4 


4.4 


Plain 


4.8 


3.7 


2.5 


2.5 


Total 


4.8 


8.9 


6.9 


6.9 


Valve (assembled) 


0.8 


0.8 


0.93 


0.93 


Loads from maximum explosion 










pressure, pounds per square 










inch: 










Assumed maximum explosion 










pressure 


510 


510 


465 


527 


Total load on piston head, pounds 


10,000 


10,000 


8,240 


9,320 


Load on wristpin 


4,000 


4,000 


3,290 


3,730 


Load on crankpin 


1,872 


1,872 


1,548 


1,751 


Load on main bearings 


(1-5)1,173 


(1-5)1,173 


(1, 2, 4)1,066 


(1, 2, 3, 5, 6) 
1,207 




(6) 1,393 


(6)1,393 


(3)694 


(4)682 




(7)364 


(7)364 


(5)402 


(7)455 



ENGINE DIMENSIONS AND ARRANGEMENTS 



77 



Continued 











Austro- 


Basse- 




Hispano-Suiza 


Benz 


Maybach 


Daimler 


Selve 


Mercedes 


180 


300 


200 


300 


200 


270 


180 


Magneto 


Magneto 


Magneto 


Magneto 


Magneto 


Magneto 


Magneto 


Splitdorf 


Splitdorf 


Bosch 


Bosch 


Bosch 


Bosch 


Bosch 


26° 


26° 


32° 


38° 


40° 




30° 


-4° 


-4° 


-5° 










CS 


CS 


1.5 X CS 


1.5 X CS 


1.5 X CS 


1.5 X CS 


1.5 X CS 


cs 


CS 












1L-4R-3L- 


1L-4R-3L- 


1-5-3-6-2-4 


1-5-3-6-2-4 


1-5-3-6-2-4 


1-5-3-6-2-4 


1-5-3-6-2-4 


2R-4L-1R- 


2R-4L-1R- 












2I^3R 


2L-3R 












Counter 


Counter 


Counter 


Counter 


Counter 




Counter 


Clock 


clock 


clock 


clock 


clock 




clock 


2 


2 


2 


2 


2 


2 


2 


Side of 


Side of 


Side of 


Cyl. head 


Side of 


Side of 


Side of 


head 


head 


head 




head 


head 


head 


18.0 


18.0 


18.0 










1.5 


1.5 


1.5 










0.021 


0.021 


0.012 










1.2 X CS 


0.933 XCS 


0.737 X CS 


0.5 XCS 


0.667 XCS 






1.2 X CS 


1.2 XCS. 


1.5 X CS 


2.0 XCS 


1.894 XCS 




1.5 X CS 


3.6 


5.95 


5.0 


14.05 


4.1 


6.19 




1.0 


1.25 


1.4 


3.30 


1.6 


2.25 




3.1 


4.40 












1.5 


1.80 


4.4 


5.62 


3.18 


6.75 




4.1 


5.65 












2.5 


3.05 


5.8 


8.92 


4.84 


9.00 




6.6 


8.70 


5.8 


8.92 


4.84 


9.00 




0.9 


0.95 


0.60 


0.881 


0.50 


0.75 




500 


500 


550 










8,768 


11,920 


13,120 










3,360 


3,860 


3,612 










1,878 


2,250 


2,273 










(1)498 


(1)588 


(1)1,959 










(2, 3, 4)745 


(2, 3, 4) 
1,192 


(2, 3, 4, 5, 6) 
1,586 










(5)4,384 


(5)5,960 


(7)1,118 











78 



THE AIRPLANE ENGINE 






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ENGINE DIMENSIONS AND ARRANGEMENTS 



79 



The weight per cubic inch of piston displacement per stroke is 
an excellent measure of the success of the designer in keeping 
down weight. The actual horse power developed depends on 
mean effective pressure and revolutions per minute and is affected 
by factors outside the engine proper: the mean effective pressure 
is determined largely by the fuel used and the carburetor and 
manifold resistance; the revolutions per minute are limited by the 
desirable propeller speed in ungeared engines. 

Detailed dimensions of some of the most successful American 
and German engines are given in Table 4. Weights of engine 
parts are given in Table 5. 

External over-all dimensions for selected engines are given 
in Table 6. The maximum width occurs at the crankcase level 



Table 6. — Over-all Dimensions of Engines 



Type 



Name 



Horse 
power 



No. of 

cylinders 



Dimensions, inches 



Maximum 



Length 



Width 



Height 



Rotary . 



Radial 

air-cooled 



Vertical . 



45-deg. Vee. 



60-deg. Vee. . 



9,0-deg. Vee.. 



f Clerget... 

I Gnome. . . 

Le Rhone. 



j ABC Wasp 

1 ABC Dragon Fly. 



Curtiss K6 

Liberty 6 

Beardmore 

Galloway 

Hall-Scott A5.. 
Hall-Scott A7. . 
Siddeley Puma. 
Austro-Daimler. 



Liberty. 



Sunbeam Cossack 

Rolls-Royce, Eagle 8. . 
Rolls-Royce, Falcon 3 

Sunbeam Maori 

Packard 

Curtiss K12 



Sunbeam Arab L. 

Curtiss OX 

Curtiss VX 

Hispano-Suiza. . . 



130 

100 

80 

170 
320 

200 
200 
160 
240 
125 
95 
240 
200 

400 

320 
360 
220 
250 
270 
400 

200 

90 

160 

180 



36.22 | 40.15 
24.75 37.5 
36.1 37.25 



35.7 
42.1 

63.0 

67.25 

57.08 

67.32 

63.38 

57.0 

69.88 

68.9 

69.1 

61.81 
75.98 
72.04 
55.41 
62.25 
68.3 

43.5 
50.0 
67.38 
51.3 



42.2 
48.5 

22.37 

19.0 

19.92 

20.07 

18.5 

18.5 

24.09 

22.4 

26.8 

37.79 

42.52 

37.24 

35.46 

27.125 

27.9 

31.7 
30.0 
45.5 
33.5 



40.15 

37.5 

37.25 

42.2 

47.7 

39.25 

43.0 

31.88 

42.91 

41.25 

39.5 

43.62 

45.3 

43.0 

38.89 

48.03 

42.0 

33.85 

34.75 

40.1 

35.5 
27.0 
35.92 
32.7 



80 THE AIRPLANE ENGINE 

in vertical engines and at the cylinder tops in Vee engines. As 
the tops of the cylinders are often above the fuselage, the width 
of the fuselage is not necessarily controlled by the maximum 
width of the engine. 

AMERICAN ENGINES 

Liberty Engine. — The Liberty engine is constructed either as a 
six-cylinder vertical, or a 12-cylinder Vee with an included angle 
of 45 deg. It has built-up steel cylinders, overhead valves and 
camshaft, and battery ignition. The cylinder units are the same 
in both constructions. Detailed dimensions are given in Table 
4; weights are given in Table 5. Longitudinal and transverse 
sections of the 12-cylinder engine are shown in Figs. 46 and 47. 
The performance of this engine is shown in Fig. 48; the full 
throttle curves are from tests with a dynamometer load and are 
carried up to 2,000 r.p.m.; the propeller load curves are from 
tests of the engine equipped with its proper propeller and 
mounted on a torque stand. With the propeller used the engine 
runs at about 1,700 r.p.m. at full throttle at ground level. Maxi- 
mum power is obtained at about 1,850 r.p.m. and maximum 
economy at about 1,800 r.p.m. The brake mean effective pres- 
sure, mechanical efficiency, and manifold depression are shown 
in Figs. 27, 14, and 17 respectively. The gear trains for driving 
the camshafts and various accessories are shown in Figs. 49 and 50 
for six- and 12-cylinder constructions respectively. Some of the 
details of this engine are discussed under the appropriate headings 
in Chapters VI and VII. 

At the rated speed of 1,700 r.p.m. the six-cylinder engine 
develops 232 h.p., and the 12-cylinder engine 425 h.p. The 
diminution in horse power per cylinder results from the lower 
mean effective pressure (see Fig. 27), which apparently results 
from lower volumetric efficiency. The weight per horse power 
falls from 2.45 lb. for the six-cylinder to 1.99 lb. for the 12- 
cylinder engine. 

Packard. — This engine is built both as an eight- and 12-cylinder 
engine, with an included Vee angle in both cases of 60 deg. 
It is very similar to the Liberty engine in its general features but 
has smaller bore and stroke, a different method of cylinder 
construction (see Fig. 92) and an underneath carburetor with 
induction pipes through the crankcase. Detailed dimensions 
are given in Table 4; weights are given in Table 5. The per- 



ENGINE DIMENSIONS AND ARRANGEMENTS 



81 




82 



THE AIRPLANE ENGINE 



formance of the 12-cylinder engine is shown in Fig. 50. With 
the propeller used in the tests the engine speed is 1,600 r.p.m. 
at full throttle; maximum power is at about 2,400 r.p.m., and 
maximum economy at about 1,800 r.p.m. The brake mean 
effective pressure and mechanical efficiency are shown in Figs. 
27 and 14 respectively. 

At the rated speed of 1,600 r.p.m. the eight-cylinder engine 
develops 192 h.p., and the 12-cylinder engine 280 h.p. The 




Fig. 47. — Transverse section of Liberty-12 engine. 

horse power per cylinder remains constant. The weight per 
horse power falls from 2.82 lb. for the eight-cylinder engine to 
2.62 lb. for the 12-cylinder engine. 

Hispano-Suiza. — This engine is built in several sizes as an 
eight-cylinder Vee with included angle of 90 deg. In this 
country two cylinder sizes are built. The design is characterized 
by steel cylinder sleeves screwed into aluminum water jackets 
cast in blocks of four, by overhead valves and camshafts, and by 



ENGINE DIMENSIONS AND ARRANGEMENTS 



83 













































450 


































































<7 


--' 














400 






















^ 


S 


/ 




























u 


^ 








r 
















350 












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300 


























































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200 










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150 








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)00 


















































































0.65 














































\ 




































0.60 








N 


& 










































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Fuel Consumption 








0.55 






















































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0.50 






















V 






Fi/ll 


n 


rtrf 


1^ 
































i 






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1 







2000 



Fig. 



1200 1400 1600 1800, 

Engine Revolutions per Minute 

48. — Performance curves of Liberty-12 engine 



£3*2: ■ *.*■»«*_ c Tt!Li&. 



drive,? crank- ^ 
shaft speed 



2 crankshaft 
-£ speed 



k shaft speech 



Generator, 
crankshaft speed 




1% crankshaft 

v' speed Tachometer' 
drive,^ crank- - 
shaft speed 

7 Crankshaft 



Water pump 
>2 crankshaft si. 

Oil pump, 
>2 crankshaft 



speed. 



m 



Water pump, 
I? crankshaft speed ft? 




lj> crankshaft 
speed 

Generator, 

li crankshaft 

speed 



Crankshaft 
gear 



speed 
Side View Liberty 6 End View Liberty |2 

Fig. 49. Fig. 50. 

Gear trains of Liberty engines. 



84 



THE AIRPLANE ENGINE 



the marine type of connecting rods. Magneto ignition is used. 
The models are known as E (180 h.p.) and H (300 h.p.). The E 
engine with a lower compression ratio (4.72) and running at 
1,450 r.p.m. develops 150 h.p. and is known as model I. 

Detailed dimensions of models E and H are given in Table 4; 
weights are given in Table 5. Transverse and longitudinal 
sections of model E are shown in Figs. 51 and 52. The per- 
formance of the 300-h.p. engine is shown in Fig. 53. With the 
propeller used in the tests the engine speed is 1,800 r.p.m. at full 
throttle; the maximum power is at about 2,300 r.p.m. and maxi- 




Fig. 51.; — Transverse section of Hispano-Suiza 180. 

mum economy at about 1,900 r.p.m. The brake mean effective 
pressure, mechanical efficiency, and manifold depression are 
shown in Figs. 27, 14 and 17, respectively. The gear trains 
for driving the camshafts and various accessories are shown in 
Fig. 54. Details of this engine are described later. 

At the rated speed of 1,800 r.p.m. the smaller engine develops 
187 h.p., the larger engine 327 h.p. The weight per horse power 
falls from 2.57 lb. for the smaller to 1.94 lb. for the larger engine. 

Wright Engine. — This engine is a slightly modified Hispano- 
Suiza. The cylinder jacket is a little lower, cylinder heads 



ENGINE DIMENSIONS AND ARRANGEMENTS 



85 




86 



THE AIRPLANE ENGINE 



thicker and the lubrication system is altered. These modifica- 
tions make for greater durability. 



380 



340 



300 



260 



220 



I 80 



40 



0.7 



0.6 





























































a 






















My 






















y 




















9. 


7 






















/ 




i/_ 














































V 






















9>-j 
































































































5 C- 
























X 


























K 


























fe 


























^// throW<L 









0.5 

1400 1600 1800 2000 2200 2400 

Engine Revolutions per Minute 

Fig. 53. — Performance curves of Hispano-Suiza 300. 



Camshaft, 
| crankshaft 
speed 







Pressure oil pump, 
| crankshaft speed 



Fig. 54. — Gear train of Hispano-Suiza 180. 



Curtiss. — The most recent Curtiss models are the K-6 and 
K-12; earlier models include the OX, VX, V-2. General data on 
these engines are given in Table 3. The K-6 is a six-cylinder 



ENGINE DIMENSIONS AND ARRANGEMENTS 



87 




88 



THE AIRPLANE ENGINE 



vertical engine; the K-12 is a 12-cylinder 60-deg. Vee engine with 
the same unit cylinder. These engines are characterized by the 
following features: cylinder blocks and top half of crankcase in 
one aluminum casting; cylinder heads in a separate aluminum 
casting bolted to the cylinder block; steel cylinder liners screwed 
into the cylinder heads; packed watertight joint between cylinder 
liners and aluminum blocks; separate overhead camshafts for the 




Transverse section of Curtiss K-12. 



inlet and the exhaust valves; two inlet and two exhaust valves 
per cylinder; sliding cam follower; four-bearing crankshaft with 
counterweights; crankshaft bearings supported by partition walls 
of upper half of crankcase; ribbed pistons. The K-6 is direct 
drive; the K-12 has a herring-bone reduction gear with a ratio 
of 5:3, and an articulated connecting rod. 

Longitudinal and transverse sections of the K-12 are shown 



ENGINE DIMENSIONS AND ARRANGEMENTS 



89 



in Figs. 55 and 56 respectively; a transverse section of K-6 is 
shown in Fig. 57. 

At the rated speed of 2,250 r.p.m. (propeller speed 1,350 r.p.m.) 
the K-12 develops 385 h.p., corresponding to a mean effective 
pressure of 119 lb. per square inch; at 2,550 r.p.m. the mean effec- 



FlG 




Transverse section of Curtiss K-6. 



tive pressure is 113 lb. per square inch and the engine develops 
415 h.p. The weight dry without exhaust manifold is 665 lb., 
giving a weight of 1.73 lb. per horse power at rated speed. The 
K-6, weighing 420 lb., develops approximately 200 h.p., which 
gives a weight of 2.1 lb. per horse power. Performance curves 
of the K-12 are given in Fig. 58. 



90 



THE AIRPLANE ENGINE 



The OX model is an eight-cylinder 90-deg. Vee. It has 
separate cast-iron cylinders with brazed monel-metal jackets; 
the cylinders in the two rows are staggered with reference to one 
another so as to permit the use of side-by-side connecting rods 
on each of the four crankpins. The valves are inclined to the 
cylinder axis and are operated from a camshaft inside the crank- 
case. A push rod operates the exhaust valve through a rocker 
arm; a pull rod operates the inlet valve. The piston is of alu- 
minum. The single carburetor is below the crankcase and has 
long intake pipes leading to the inlet manifold. 

Longitudinal and transverse sections of this engine are shown 
in Figs. 59 and 60. Performance curves of the OX-5, as publish- 
ed by the builder, are given in Fig. 61. 



420 



400 



I 380 



o 360 



E 340 



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100 



1100 1200 1300 1400 1500 1600 

Revolutions per Minute of Propeller Shaft 

Fig. 58. — Performance curves of Curtiss K-12. ! 



The V-2 model is similar in general arrangement to the OX 
but is of larger size. The cylinders are of steel with welded 
monel-metal jackets; the valve stems are parallel to the cylinder 
axes. Two carburetors are used and are located at the level of 
the bottom of the crankcase. When used for low level flying 
(up to 6,000 ft.) an aluminum liner is used between the cylinders 
and the crankcase; for high flights these shims may be taken out 
and the compression ratio thereby increased. A transverse view 
of this engine is shown in Fig. 62. 

Hall-Scott.— General data on several models built by this 
company are given in Table 3. The latest model is the L-6; it 
has the same bore and stroke as the A-5, A-7 and A-8 models and 
also the Liberty engine. Longitudinal and transverse sections of 



ENGINE DIMENSIONS AND ARRANGEMENTS 



91 




92 



THE AIRPLANE ENGINE 




Fig. 60. — Transverse section of Curtiss OXX-3. 



110 

100 

90 

80 

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D. 

250© 

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150 
100 
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J00 



2100 



1100 1300 1500 1700 1900 
Revolutions per Minu+e 

Fig. 61. — Performance curves of Curtiss OX-5. 



ENGINE DIMENSIONS AND ARRANGEMENTS 



93 



this engine are shown in Figs. 63 and 64. It is very similar to 
the Liberty 6 in its cylinders, valves, pistons, camshaft, connec- 
ting rod and crankshaft. The crankshaft is supported entirely 
by the upper half of the crankcase, the bearing caps being bolted 
through the crankcase by through bolts which on the upper end 
act as cylinder hold-down bolts. The piston pin floats freely in 




Fig. 62. — Transverse section of Curtiss V-2. 



both the rod and piston. Carburetion is through two carburetors 
and hot-spot water-jacketed manifolds. 

Performance curves of the four cylinder L-4, as published by the 
builder, are given in Fig. 65. 

Bugatti. — The King-Bugatti engine, built in the United 
States, is a twin-vertical 16-cylinder engine with the two crank- 
shafts geared to drive a common propeller shaft. It is a modifi- 



94 



THE AIRPLANE ENGINE 




ENGINE DIMENSIONS AND ARRANGEMENTS 



95 



cation of the French Bugatti engine. It has a number of special 
features which may be seen in the sections, Figs. 66 and 67. The 
cylinders are of iron, cast in blocks of four, with integral water 
jackets except at the sides, which are covered with cast aluminum 
plates bolted to the cast iron. There are two inlet valves and one 
, exhaust valve per cylinder. 

The crankshaft bearings are 
supported from the upper 
part of the cast aluminum 
crankcase; the caps, each of 
which supports two bearings, 
extend nearly the whole width 
of the crankcase. The eight- 
throw, nine-bearing crankshaft 
is in two halves connected at 
the center by a taper and key, 
shrunk and drawn up by a nut. 
Each half of the shaft forms 
a four-cylinder shaft with the 
throws all in one plane; the 
throws of the two sections are 
assembled at right angles. In 




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130 | 

120 % 

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110 .£ 

100 « 

90 I 



Fig. 64.- 



-Transverse section of Hall-Scott 
L-6. 



80 

1000 1200 1400 1600 18.00 
Revolu+ions per Minute 

Fig. 65. — Performance curves of Hall- 
Scott L-4. 



assembling the completed crankshafts in the crankcase they are 
placed in such relation to each other that if No. 8 throw left is on 
top dead center, No. 8 throw right will be 45 deg. past bottom 
dead center. All bearings, including the connecting rod bearings, 
are undercut and thereby shorten the over-all length of the engine 
by about 5 in. The valves are operated by a single camshaft for 



96 



THE AIRPLANE ENGINE 




ENGINE DIMENSIONS AND ARRANGEMENTS 



97 



each half of the engine. This camshaft is operated through 
a bevel gear, driven by a vertical shaft between the two cylinder 
blocks, which in turn is driven by a bevel gear on the crankshaft. 




Fig. 67. — Transverse section of Bugatti engine. 

The magneto drive (see Fig. 245) is from a bevel which is on this 
vertical shaft. The rocker arms are pivoted on rods running 
along each side of each camshaft housing. The cam follower 
is a hardened-steel roller; a smaller roller operates direct on a 



98 



THE AIRPLANE ENGINE 



cap on top of the valve stem. The hollow propeller shaft runs 
at two-thirds engine shaft speed. 

A performance curve of this engine both with dynamometer 
and propeller load is shown in Fig. 68. Maximum horse power 
is at an engine speed in excess of 2,400 r.p.m. 




1&00 1900 2000 2100 2200 
Engine Revolutions per Minute 

Fig. 68. — Performance curves of Bugatti engine. 



2400 



ENGLISH ENGINES 

Rolls-Royce. — The Rolls-Royce engines are interesting as 
probably representing the highest grade of design and manu- 
facture in any country. General dimensions of the 12-cylinder, 
60-deg. Vee "Eagle" and "Falcon" models are given in the 
following table, which shows how the power and efficiency 
of these engines have been improved by increasing the 
revolutions per minute and the compression ratio. Sectional 
views are given in Figs. 69 and 70. These engines have an 
epicy clic reduction gear concentric with the crankshaft (see p. 38 1 ) ; 
this gear is contained in a housing bolted to the front of the 
crankcase and is not shown in Fig. 69. The housing for the 
driving gears of the valve motion, magnetos, and other accessories 
is bolted to the rear of the crankcase. The upper part of the 
crankcase carries the bearings of the crankshaft; the lower part 
is merely a deep oil well. Each cylinder is fastened to the 
crankcase by four bolts. The engine is supported by arms which 



ENGINE DIMENSIONS AND ARRANGEMENTS 99 




100 THE AIRPLANE ENGINE 

are bolted to vertical surfaces at the sides of the crankcase, 
thus permitting the ready adaptation of the engine to the airplane 
without the need of special engine bearers in the fuselage. 



Fig. 70. — Transverse section of Rolls-Royce "Eagle." 

The cylinders are of steel with valve fittings welded to 
short nipples in the cylinder head and with welded jackets. The 
aluminum piston is of special design without a skirt at the middle 
third of its length (see Fig. 98, p. 135); this design transfers 
the gas pressures directly to the piston bosses and reduces the 



ENGINE DIMENSIONS AND ARRANGEMENTS 



101 



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102 



THE AIRPLANE ENGINE 



bending stress in the piston head. The piston pin is close to the 
lower edge of the piston. 

The valves are of tulip style (see Fig. 115, page 157 ), with 
bored stems. Inlet and exhaust valves are interchangeable. A 
hardened contact button is set in the end of the valve stem. The 
crankshaft is provided with balance weights of cast steel bolted 
to each crankarm; these have increased the weight of the engine 
but have reduced the pressures on the main bearings. The 
forward end of the crankshaft carries a flange to which is secured 



0.60 




Fig. 71. 



1400 1500 1600 1700 1300 1900 
Revolutions per Minute 

-Performance curves of Rolls-Royce 



Eagle." 



the internally-toothed reduction-gear wheel (see Fig. 296). A 
stud pressed into the bore of the crankshaft at the forward end 
serves as pilot for the spider of the planetary pinions. The con- 
necting rods are of the articulated type, the secondary rods 
working on pins which are clamped into lugs formed on both 
sides of the main connecting rod. 

The camshaft drive is from a worm on the crankshaft, which 
is connected to the shaft through a serrated hub joint, and a 
spring coupling clamped by means of a disc brake; this device 
protects the drive against rotary vibration. The worm drives a 
cross shaft which has bevel wheels at its ends driving the inclined 
intermediate shafts going to the camshafts. The magnetos are 
driven from the cross shaft. 



ENGINE DIMENSIONS AND ARRANGEMENTS 103 

Performance curves for the " Eagle" with various fuel valve 
settings are given in Fig. 71. 

Napier "Lion."— The Napier "Lion" is the only W or " arrow" 
type of airplane engine that is definitely successful. Details of 




this engine are shown in Figs. 72 and 73; dimensions in Table 3. 
It has three blocks, of four cylinders each, mounted on a single 
crank case, with an angle of 60 deg. between the rows. This engine 
is probably the lightest of all the successful water-cooled engines; 



104 



THE AIRPLANE ENGINE 



it weighs 1.86 lb. per horse power, dry, and 2.51 lb. per horse power 
with its jackets full of water. 

The cylinders are 5|-in. bore and 5§-in. stroke which is an 
unusually low stroke-bore ratio and makes for small over-all height 
and width. Each block of cylinders is built up and consists of 




four steel liners with sheet-steel water jackets. The cylinders of 
each block are secured to an aluminum head casting to which they 
are fastened by valve seats, which pass through the crown of each 
cylinder -and screw into the head casting. The head casting 
carries camshafts and their bearings and drives, and contains 



ENGINE DIMENSIONS AND ARRANGEMENTS 105 

the inlet and exhaust ports and passages. Each cylinder has two 
inlet and two exhaust valves, the latter being on the outside of the 
inclined blocks. The valve guides are bronze and a tight fit in 
the cylinder head. The inlet and exhaust camshafts operate the 
valves directly without intermediate rockers or plungers; each 
valve has an adjustable tappet head. One of each pair of cam- 
shafts is driven by bevel gears on the inclined and vertical shafts 
leading from the distribution gearing; the other camshaft is driven 
by spur gearing from the first shaft. The crankshaft has four 
throws and is supported on five roller bearings; the roller bearings 
at the front and rear are fitted direct to the shaft; the three 
intermediate ones have large inner races, which permit the 
bearings as a whole to be threaded over the crank webs and which 
are mounted on split bushings keyed to the shaft. The connect- 
ing rod assembly consists of a master rod and two side rods 
carried on lugs integral with the big end of the master rod. The 
pistons are unusually shallow, having a depth of only 3f in. The 
firing order is : 

Propeller End 

7 2 9 

4 11 6 
10 5 12 

1 8 3 

The propeller shaft is mounted on roller bearings with double 
ball-bearing thrust block. The ratio of propeller to engine speed 
is 29 to 44. The performance curve for this engine is given in 
Fig. 74. The brake horse power is still increasing rapidly at 
2,100 r.p.m.; the mean effective pressure is a maximum at 1,800 
r.p.m. 

Siddeley "Puma." — This engine is a typical six-cylinder 
vertical engine of about 250 h.p. at 1,400 r.p.m. Cross-section 
views are given in Figs. 8 and 9; general dimensions in Table 3. 
The cylinder construction consists of steel liners in aluminum 
heads and jackets. The heads are cast in sets of three and incor- 
porate the valve ports and head jackets. The barrel jackets 
are also cast in sets of three and are bolted to the heads. The 
steel liners are screwed cold into the heads which are heated to 
about 300°C. The water joint at the lower end of each jacket is 
made by a screwed gland compressing a rubber ring. The 
three liners have to be trued up by surface grinding after assembly 



106 



THE AIRPLANE ENGINE 



into a unit. The inlet and exhaust valve seats are of phosphor 
bronze expanded into the aluminum. The difficulties winch 
others have met in using this material at very high temperatures 
have been overcome by using an exceptionally hard-chilled alloy. 
There are two exhaust valves and one inlet valve per cylinder. 
The cooling water enters the cylinder head at two places; one is 
in direct communication with the water space; the other connects 
with an aluminum tube which runs inside the full length of the 
jacket and directs comparatively cool water on to the hottest 
places. The pistons, connecting rods and crankshaft are of 



480 



460 



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i 400 



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130 a. 

UJ 

120 s 



J500 1600 1700 1800 1900 2000 2100 
Revolutions per Minu+e 

Fig. 74. — Performance curves of 
Napier "Lion." 



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160 











120^ 
110 



900 1100 1300 1500 1700 
Revofuf ions per Minute 

Fig. 75. — Performance curves of 
Siddeley "Puma." 



conventional form. A roller bearing and double-thrust ball bear- 
ing are at the propeller end of the crankshaft; the five intermediate 
shaft bearings are supported by the upper half of the crankcase. 
The bearing caps are of aluminum reinforced with webbed plates 
of steel; the caps of the two outer journals are integral with 
the lower half of the crank case. 

The camshaft is placed off-center and directly over the exhaust 
valves. It is made of steel tubing with cams pinned in position, 
an arrangement making for easy and cheap replacement of worn 
cams. The exhaust cams act directly on adjustable tappet heads 
on the valve stems; the inlet cams act on short drop-forged rock- 
ers without rollers. The valves are trumpet-shaped and give good 
flow lines. 



ENGINE DIMENSIONS AND ARRANGEMENTS 107 

A performance curve for this engine is shown in Fig. 75; 
maximum mean effective pressure is developed at about 1,000 
r.p.m.; maximum horse power well above 1,700 r.p.m. 

ITALIAN ENGINES 

Fiat. — The Fiat 650-h.p. is one of the most powerful aircraft 
engines in use at present. It is a 12-cylinder 60-deg. Vee engine 
of unusually large cylinder dimensions, 170 by 210 mm. A 
longitudinal section of this engine is given in Fig. 76; general 
dimensions are given in Table 3. 

The cylinders are of built-up steel construction with the 
cylinder heads integral with the barrels and with welded water 
jackets. Two inlet and two exhaust valves are located in the 
head of each cylinder at an angle of 25 deg. to the cylinder axis. 
The valve stems work in phosphor-bronze bushings. Each 
pair of valves is closed by duplex springs located between the 
valve stems (Fig. 134). The camshaft for each row of cylinders 
is in two parts which are driven by bevel gears from the center of 
the length of the engine. A tubular lay shaft, on ball bearings, 
is mounted in the center of the crankcase in the top of the Vee and 
is driven by spur gearing from the rear end of the crankshaft; 
it extends only to the center of the engine where it is fitted with a 
bevel gear driving the two inclined shafts which operate the 
camshafts. The connecting rods are forked type, the forked 
rods being fitted with bronze bearing shells with white-metal 
liners; the center rod has a case-hardened steel finer running on 
the outside of the bronze shell of the forked rod. The crankshaft 
is of conventional design; the main bearings are held between the 
two halves of the crankcase, that is, the lower half of the crank- 
case has the bottom halves of the bearing housings cast integrally 
with it. The water pump is below the engine at the middle of its 
length; the oil pumps are below the two ends of the lower crank- 
case. Water and oil pumps are driven from a lay shaft, at the 
bottom of the lower crankcase, which receives its motion from the 
crankshaft through spur gears at the rear of the engine. The 
induction manifolds are of sheet copper with steel flanges brazed 
at the joints. All distribution gears and driving shafts are 
mounted on ball bearings. There are four spark plugs per 
cylinder receiving current from four 12-cylinder magnetos. The 
normal output of the engine is 650 h.p. at 1,500 r.p.m.; the maxi- 
mum is 720 b.h.p. at 1,700 r.p.m. 



108 



THE AIRPLANE ENGINE 




ENGINE DIMENSIONS AND ARRANGEMENTS 109 




110 



THE AIRPLANE ENGINE 



GERMAN ENGINES 

Benz. — The Benz 230-h.p. engine has six vertical water-cooled 
cylinders 140 by 190 mm. Longitudinal and transverse sections 
of this engine are given in Figs. 77 and 78; dimensions in Table 4. 




Fig. 78. — Transverse section of Benz 230. 

The cylinders are of cast iron with pressed-steel water jackets 
around the barrels; they are bolted to the crankcase by long 
studs which pass through the upper crankcase and are screwed 
into the bottom halves of the crank bearing housings which are 
cast integral with the lower half of the crankcase. The^pistons 



ENGINE DIMENSIONS AND ARRANGEMENTS 



111 



are of cast iron with the heads supported by conical steel forgings 
riveted and welded to the piston crown and bearing on the piston 
pins through slots cut in the small ends of the connecting rods; 
by this construction the greater part of the force of the explosion 
is transmitted directly to the connecting rod. The connecting 
rods are tubular with internal oil pipes. 

Two inlet and two exhaust valves are arranged vertically on 
each cylinder head and are operated through rockers mounted on 
ball bearings. The rockers actuate the valves through rollers 
mounted on eccentric bolts which permit a fine adjustment for the 

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800 1000 1200 1400 

Revolutions per Minu+e 

Fig. 79. — Performance curves of Benz 230. 



tappet clearance. There are separate camshafts for the two 
sets of valves; these shafts are of steel tubing and are inside the 
crankcase. They are driven by spur gears from the crankshaft 
through an intermediate gear. The push rods have hemi- 
spherical ends at the bottoms which work in steel cups inside the 
hollow tappets. The tappets are slightly offset from the cam- 
shaft centers and carry steel rollers. 

The upper half of the crankcase has transverse air passages 
cast in the webs which serve to cool the crankcase and to supply 
warm air to the carburetors. The lower half is cooled by trans- 
verse aluminum tubes, the air beingjed into one side by a large 
louvered cowl. 

Performance data on this engine are given in Fig. 79. 



112 



THE AIRPLANE ENGINE 




ENGINE DIMENSIONS AND ARRANGEMENTS 



113 



Maybach. — The Maybach 300-h.p. engine has six vertical 
cylinders built up of steel liners screwed into cast-iron cylinder 
heads and with forged steel jackets screwed to the cylinder heads. 
Sectional views of the engine are given in Figs. 80 and 81; general 
dimensions in Table 4. The compression ratio of this engine, 
5.94 : 1, is unusually high. The pistons are of cast iron. Con- 
necting rods are of square section, bored out; the small end has a 
cast-iron floating bushing. 

There are two inlet and 
two exhaust valves work- 
ing vertically in cast-iron 
guides in each cylinder 
head and operated by 
rocker levers mounted on 
roller bearings, each pair 
of valves being operated 
by a single tappet rod 




300 

2&0 

260 

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I ISO 

160 

140 
120 





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1000 1200 1400 1600 1800 
Revolutions per Minute 

Fig. 82. — Performance curve 
of Maybach 300. 



from one of the camshafts in the crankcase. The bottom halves 
of the crankshaft bearings are very deep and are bolted to the top 
half of the crankcase. The lower hah of the crankcase supports 
the oil pumps and carries detachable oil pumps. 

The performance curves, Fig. 82, show maximum mean effec- 
tive pressure at 1,300 r.p.m., maximum power at 1,550 r.p.m., 
and maximum economy at 1,300 r.p.m. 

8 



CHAPTER V 



MATERIALS 



The structural materials available for airplane engine parts 
are in five classes, steels, carbon and alloy and either forged 
or cast, cast iron, aluminum alloys, bronze, and bearing metals 
or babbitts. The choice of metal for any given part is determined 
principally by (1) the suitability of the metal for meeting the 
stresses and general operating conditions to which the part is 
to be subjected, (2) the weight, and (3) the machinability 
of the material. It is obvious that condition (1) must always 
be met, but it is often possible, especially with lightly stressed 
members, to meet that condition with a number of different 
materials; for example, the crankcase may be of cast iron or 
aluminum. The selection as between these materials will 
usually be on the basis of weight although machinability 
or cost may determine the final choice. From the point of 
view of tensile strength alone the important quantity is not 
strength per unit of cross-section area but per unit of weight. 
On this basis the metals have the following properties i 1 



Weight per 
cubic foot 



Tensile 

strength, 

pounds per 

square inch 



Tensile 
strength 4- 
weight per 
cubic foot 



Aluminum alloy, cast. . 
Aluminum alloy, forged 

Cast iron 

Gun metal (bronze) 

Malleable iron 

Cast steel 

Mild steel 

High tensile steel 

Nickel-chrome steel 



170 
170 

480 
500 
480 
480 
480 
480 
480 



27,000 
60,000 
20,000 
31,000 
40,000 
60,000 
60,000 
100,000 
135,000 



159 

350 

42 

62 

83 

125 

125 

208 

280 



The advantage of aluminum over cast iron for castings is not 
only in the great reduction in weight but also in the much greater 
1 Pomeroy, Jour. Soc. Aut. Eng., Jan., 1920. 

114 



MATERIALS 115 

facility for machining. The substitution of a lighter material in 
a stressed member may result in considerable saving in weight 
even if the strength-weight ratio of the lighter material is not so 
favorable as in the heavier material. With castings this results 
from the fact that they cannot be reduced below certain thick- 
nesses, which are often in excess of strength requirements, because 
of foundry considerations. For machined members, such as 
connecting rods, it is also undesirable to go below certain thick- 
nesses, and additional material has to be left to avoid high inten- 
sity of stress at fillets where the cross-section is changing rapidly. 
By using a lighter metal it may be possible to load all parts of the 
member up to the allowable working stress, and thereby to 
diminish the weight. 

An important design factor affecting the choice of metal is the 
fact that the strength of members subjected to bending or torsion 
is proportional to the cube of the cross-sectional linear dimen- 
sions, while the weight is proportional to the square of the linear 
dimensions. Thus, comparing two beams or shafts of the same 
material and length and with similar cross-sections but with 
linear dimensions of 1 and 1.41 respectively, the second will be 
twice as heavy as the first, and (1.41) 3 = 2.8 times as strong. 
Stiffness is also an important consideration, and as this varies as 
the fourth power of the linear cross-section dimensions, the 
second beam or shaft in the above example would be (1.41) 4 = 4 
times as strong as the first one. If, in the second case, the 
allowable stress on the material is only half of that permissible 
in the first case, but the strength-weight ratio is unchanged, 
the weight will be unchanged, the strength increased 1.4 times 
and the stiffness doubled. For the same strength for a beam 
or shaft, with a stress in the second case one-half that in the first 
case, and constant strength-weight ratio, the relative linear 
dimensions would be 1 to v"2, or 1 to 1-26, and the relative 
weights 1 to (1.26) 2 , or 1 to 1.588. 

Materials for Special Parts. — Cylinder liners are nearly always 
of steel. It is not necessary to use a steel of high tensile strength 
since the liners cannot be machined with safety down to the 
dimension which will stress the material fully. The use of an 
aluminum alloy may be desirable where short life is expected (as 
in military use), but its resistance to abrasion is low although its 
heat conduction is much superior to that of steel. 

Crankshafts are dimensioned for stiffness as well as for strength. 



116 THE AIRPLANE ENGINE 

Stiffness for given dimensions depends only on the modulus of 
elasticity and not at all upon tensile strength. As the modulus of 
elasticity is practically constant for all steels, there is no advan- 
tage, so far as stiffness is concerned, in using a steel of very 
high tensile strength for shafts. 

Connecting rods must be made as light as possible to keep down 
the unbalanced inertia forces. As they have to be machined all 
over and reduced to very thin sections it is important that the 
material should be readily machined and free from flaws. Forged 
aluminum would be excellent for this purpose when available. 

Piston materials are considered on p. 132. Valve materials 
are selected on other than strength considerations (see p. 159). 

Steels. — Steels for airplane engine use should have the follow- 
ing fundamental properties in as great a degree as possible: (1) 
high strength, (2) high toughness, (3) great durability, (4) 
soundness, (5) ease of machining, (6) ease of heat treatment, (7) 
constancy of properties, (8) homogeneity. The steels which are 
available are endless in number but divide themselves into certain 
types. Those which are of most importance for airplane engine 
construction are listed in Table 7, 1 which gives their composition, 
heat treatment and physical properties. Other special steels 
for valves are discussed on page 160. The 0.45 C steel is an 
excellent general-utility steel with good mechanical strength, 
high ductility and easily machined. The 0.9 — 1.0 C steel 
gives good service in resisting abrasion, is readily produced in 
sheet form and does not need hardening and tempering. The 
case-hardening steels are of importance for those parts which 
require local surface hardening combined with general toughness; 
the surface hardness of the 0.1 C and the 5 per cent Ni are about 
the same as shown by the Brinell test but the strength of the 
nickel steel is much higher. The air-hardening Ni-Cr steel 
shows the very high Brinell number of 477; its use is principally 
for gears. The 3 per cent Ni steel combines high tensile strength 
and high ductility; the addition of chromium increases the 
hardness. The chrome-vanadium steel has shown fewer failures 
in practice than most of the alloy steels, probably as a result of 
absence of air-hardening characteristics. The high-chromium 
or " stainless' ? steel has not only its non-rusting qualities but its 
high mechanical properties to recommend it. At present its cost 
is high. 

1 Selected from Hatfield, The Automobile Engineer, June, 1920. 



MATERIALS 



117 



a 

1 

O 

u ■ 
xi 
o 


CO CD 

00 CM O CO CM 00 O 

NNHOOWN 

o" d o" © d cm d 


Oil 
quenched 

950°C. 
Tempered 

250°C. 


212,000 

240,000 

7 

29.5 

444 
64 


73 T3 

_ Js d » * d 

rS o o -poo 
O PI © * a o 

(1) lO ^ CO lO 

3 OJ r" 3 O 


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O © CM 

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00 rH 


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it 


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T-J T3 
5 . CO . 

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w co »o , io 

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© O CM © iO -tf 
O © CM 

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N OMMMOH . 
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13 t3 

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ft d O S B O 

u S>o^ So 


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co 

S3* 
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CO W i-H O O • © • 

d d d d d -cm • 


13 13 

m « £. CO 

^ ^ o & .* o 

O pi o « pi © 

0> lO [> CO CM 
pl 00 I 5 " pi © 


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O O CM • CM CO 
O © CO CM 

CO 
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co 

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D* Eh 


© O O 00 © rj< 

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5 per cent 

Ni case 
hardening 


c 


00 © • 

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CM rH O O -00 • 


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a 
c 

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55 C 
O .£ 

a *■ 

u 

o 


Manganese, per cent 

Silicon, per cent 

Sulphur, per cent 

Phosphorus, per cent 

Chromium, per cent 

Nickel, per cent 

Vanadium, per cent 


PI 
_o 

T3 

CI 

o 
O 


Physical properties: 

Yield point, pounds per 

square inch 

Ultimate strength, pounds 


. PI 

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. o 

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1 

a. 





118 



THE AIRPLANE ENGINE 



The materials that have been employed for the various parts 
of airplane engines are given in Table 8, 1 which also states the 
nature of the stresses to which the part is subjected and the inten- 
sity of the service which it has to perform. In addition, there is 
given in the last column the material recommended by Doctor 
Hatfield on the basis of an unusually wide experience in investi- 
gating the physical properties of the steels and the causes of 
failure of the parts of airplane engines. There is naturally 
much difference of opinion among engineers as to the best 
material to use for several of these parts, and practice is by no 
means standardized. 

Aluminum Alloys are largely used in airplane engines on 
account of their light weight. The specific gravity of pure alumi- 















/ 




. c4UU 












/ 




fc 2200 
















~ 2000 




0$ 


X 










oo 1800 






*<./ 










£ 1600 




y/ 










■55 1400 
P 1200 

















D 

CD 4 

























x& 








& 






^x 








\ 



























2 



10 12 14- 



14 



Metal Alloyed with Aluminum, 
Per Cent. 

Fig. 83. — Tensile strength of 
aluminum alloys. 



2 4 <b 8 10 12 
Metal Alloyed with Aluminum, 
Percent 

Fig. 84. — Ductility of aluminum 
alloys. 



num is 2.56; of the common alloys about 2.8; of steel about 7.8. 
The important alloys are those with copper and with zinc. The 
effect of the presence of these substances on the tensile strength 

Table 8. — Materials for Engine Parts 
Glossary of Terms used. 



A. H. Ni. Cr. 


= Air-hardening Nickel-chro- 


M. C. I. 


= Malleable Cast Iron. 




mium Steel. 


M. C. Steel 


= Medium-carbon Steel. 


Al. 


= Aluminium Alloy. 


M. S. 


= Mild Steel (0.15 carbon). 


C. H. C. 


= Case-hardening Carbon 


Ni. 


= Nickel Steel (3 per cent). 




Steel. 


Ni. C. H. 


= Nickel Case-hardening Steel. 


C. I. 


= Cast Iron. 




(5 per cent Ni.). 


C. Steel 


= Carbon Steel. 


Ni. Cr. 


= Nickel-chromium steel (3 per 


Cr. Van. 


= Chromium Vanadium Steel. 




cent). 


H. C. Steel 


= High-carbon Steel. 


Ph. Br 


= Phosphor Bronze. 


H. S. 


= 12 to 14 per cent tungsten 


Si. Cr. 


= Silicon Chromium Steel. 




High-speed Steel. 


Steel C. 


= Steel Casting. 


H. T. 


= High-tensile Steel. 


T. Steel 


= Tungsten Steel. 


1 Hatfield, loc. cit. 







MATERIALS 



119 



Table 8 (Continued) 



Part 



Nature of 

chief stresses 

in service 



Intensity 

of 

service 



Materials which 

have been or 

are used 



Materials 
recommended 



Cylinder . 



High temperature, 
tension and abra- 
sion. 



Heavy or Al.; C. I.; Steel 
medium. C; 100,000 lb. 
C. Steel; NL; 3 
per cent Ni. Cr. ; 
Steel Mixture. 



Al.; 80,000 lb. 
Steel. 



Cylinder holding- 
down bolts. 



Tension and bend- Medium, 
ing. 



j Medium or M. 
S.: Ni. 



3 per cent Ni. 



Cylinder liners 


High temperature 
and abrasion. 


Heavy. 


C. I.; Steel C; 
Forged Steel. 


100,000 lb. C. 
Steel. 


Spark-plug body. . . 


High temperature. Medium. 


Brass; C. I.; M. 
S.; Stainless. 


Stainless. 


Spark-plug 
trode. 


elec- 


Very high tempera- Heavy, 
ture. 


T. Steel; Ni. 


Nichrome Alloy 
or Stainless. 


Valve cages 


High temperature, Medium. 
various slight 
stresses. 


C. I. 


M. C. I. 


Valve-rocker 
era. 


roll- 


Abrasion. 


Light to 
medium. 


C. H. C; Ni. Cr. 


5 per cent Ni. 
C. H. C. 



Valves . 



High temperature, Heavy, 
tension, shock and 
abrasion. 



H. S. Steels; 25 
per cent Ni. ; 
Stainless Steels; 
Ni. Cr.; 3 per 
cent Ni.; H. C. 
Steel. 



Stainless. 



Valve guides . 



High temperature 
and abrasion. 



Medium. 



AL; C. I.; C. 

Steel; T. Steel. 



Stainless. 



Valve seats. 



Abrasion, shock 
and high tem- 
perature. 



Heavy. C. I.; M. C. 

i Steel; Ni. Cr.; 
Ph. Br. 



80,000 lb. C. 
Steel. 



Valve springs. 



Torsion and bend- 
I ing. 



Light. 



Cr. Van. Steel; 
Si. Cr. Steel; 
H. C. Steel (oil 
hardened). 



Cr. Van. Steel. 



Valve rockers . 



Bending, shock and Medium, 
abrasion. 



Ni. Cr.; C. H. 
Ni.; Bronze. 



5 per cent Ni. 
C. H.; 3 per cent 

Ni. 



Valve-rocker bear- 
ing. 



Abrasion and com- 
pression. 



Light. 



Ph. Br. and 

White Metal. 



Ph. Br. 



Water jacket ' Very slight. 



Light. 



Al.; Copper; 
Pressed Steel; 
Sheet Steel. 



Steel Sheet. 



120 



THE AIRPLANE ENGINE 
Table 8 (Continued) 



Part 


Nature of 

chief stresses 

in service 


Intensity 

of 

service 


Materials which 

have been or 

are used 


Materials 
recommended 


Connecting rod .... 


Compression, ten- 
sion, bending and 
shock. 


Heavy. 


A. H. Ni. Cr.; 
Cr. Van. Steel; 
Ni. Cr.; Ni. 


Ni. Cr.; A. H. Ni 
Cr. 


Connecting-rod, 
big-end bearing. . . 


Abrasion and com- 
pression. 


Heavy. 


White Metal; Ph. 
Br. 




Connecting-rod, 
little-end bearing. . 


Abrasion and com- 
pression. 


Medium. 


Ph. Br.; Gun 
Metal. 




Connectin g-rod 
bolts. 


Tension and bend- 
ing. 


Medium. 


Ni.; Ni. Cr.; Cr. 
Van. Steel. 


3 per cent Ni. 




Shear, bending and 
abrasion. 


Heavy. 


Ni.; 160,000 to 
180,000 lb. C; 
C. H. C. Steel; 
C. Steel; Ni.; 
Ni. Cr.;T. Steel. 


5 per cent Ni. 
C. H. 






Temperature, bend- 
ing and other 
stresses. 


Heavy. 


Al.; C. I.; Steel 
Drawn or Pres- 
sed; Ni. Cr. 


AL; 80,000 lb. C. 
Steel. 




Piston ring 


High temperature, 
bending and 
abrasion. 


Heavy. 


C. I.;H. S. Steel. 


C.I. 


Bevel-gearshaft for 
overhead c a m- 

shaft. 


Torsion and bend- 
ing. 


Medium. 


C. Steel; Ni.; 
Cr. Van. Steel. 


5 per cent Ni. C. 
H. 




Bending and vari- 
ous slight stresses. 


Light. 


C. Steel; Al.; 
Ni. 


Al. 






Crankshaft 


Torsion, bending, 
and shock. 


Heavy. 


Cr. Van. Steel; 
Ni. Cr.; Ni. 


Ni. Cr. 


Crankshaft journal 
bearing. 


Abrasion and com- 
pression. 


Medium. 


Ph. Br.; White 
Metal. 




Propeller reducing 
gears. 


Shear and bend- 
ing, abrasion and 
shock. 


Medium. 


Ni. C. H.; Ni.; 
Ni. Cr.;Cr. Van. 
Steel; A. H. Ni. 
Cr.; C. H. C. 


A. H. Ni. Cr. 




Compression and 
abrasion. 


Heavy. 


Ni. C. H.; C. H. 
Cr. 






H. 


Cam housings 


Negligible. 


Light. 


Al. 


Al. 


Camshaft bearing. . 


Abrasion and com- 
pression. 


Medium. 


White Metal; Ph. 
Br. 


White Metal; Ph. 
Br.; Gun Metal. 




Torsion and abra- 
sion. 


Heavy. 


Ni. Cr. C. H; 
Ni.; H. C. Cr. 
Steel. 


5 per cent Ni. C. 
H. 






Compression and 
abrasion. 


Heavy. 


C. H. C. hard- 
ened on wear- 
ing surface; Ni. 
Cr. 


5 per cent Ni. 




C. H. 



MATERIALS 121 

of cast-aluminum alloy is shown in Fig. 83. It will be seen that 
copper increases the strength up to about 9 per cent, with 7 to 
8 per cent of copper there is obtained a tough alloy of a tensile 
strength of 20,000 lb. per square inch. With zinc the tensile 
strength increases up to about 35 per cent; with this alloy the 
tensile strength reaches 50,000 lb. per square inch but it is very 
brittle and has a specific gravity of 3.3. The ductilities of the 
two types of alloy are shown in Fig. 84. With both copper and 
zinc present, higher tensile strength combined with fair ductility 
can be obtained; for example, with 2.75 per cent Cu and 7 to 
8 per cent Zn a tensile strength of 28,000 lb. per square inch and 
a ductility of 8 per cent. This alloy falls off rapidly in tensile 
strength with increase in temperature; at 570°F. the strength is 
9,500 lb. per square inch. A small addition of manganese to a 
copper aluminum alloy increases the strength and maintains it 
better with increase of temperature. 

Forged-aluminum alloy is best represented by Duralumin, 
whose composition is 93.2 to 95.5 per cent aluminum, 0.5 per cent 
magnesium, 3.5 to 5.5 per cent copper and 0.5 to 0.8 per cent 
manganese. This material can be made into plates and tubes. 
The tensile strength is about 60,000 lb. per square inch but 
can be increased by rolling to about 75,000 lb. per square inch 
though with loss of ductility; the elongation of the alloy is 15 to 
20 per cent. At 460°F. the tensile strength is halved. Forged- 
aluminum is a good bearing metal. 

Both cast- and forged-aluminum alloy have a modulus of 
elasticity of about 10,000,000 lb. per square inch or about one- 
third that of steel. The stiffness of a plate structure is propor- 
tional to the modulus of elasticity and to the cube of its thickness. 
An aluminum plate of the same weight as a steel plate would be 
nearly three times as thick; its stiffness would be about eight 
times as great as that of the steel plate. 



CHAPTER VI 
ENGINE DETAILS 

Cylinders. — There is considerable variety in the design of 
modern water-cooled airplane-engine cylinders. In one impor- 
tant respect they are all in accord, namely, in the adoption of 
overhead valves, in which they differ from the common auto- 
mobile engine. They differ further from the automobile engine 
in that the cast-iron block construction is seldom used. Both 
of these changes have been made, primarily, in order to reduce 
the weight of the engine. 

Airplane-engine cylinders are formed either singly, or in 
blocks of two, three or four. The single cylinder is flexible and 
permits freedom of movement of the cylinder without putting 
strains on other parts of the engine. As the engine is not held 
rigidly it is desirable to give all its parts a maximum of freedom. 
The single-cylinder construction is used in the majority of existing 
engines and is always employed in those engines whose cylinders 
are all steel. Examples of this construction are the Liberty, 
Packard, Hall-Scott L, Curtiss K, Rolls-Royce, Napier, Renault, 
Lorraine-Dietrich, Fiat, Mercedes and Austro-Daimler engines. 
It is necessarily employed also in radial and rotary engines (see 
Chapter VIII). The block arrangement has a common jacket 
around the cylinders which may result in some slight decrease 
in weight of the jacket itself but will usually increase the weight 
of water in the jackets and give inferior water circulation. The 
major advantage of employing the block construction is that it 
permits the cylinders to be more closely spaced and tnereby 
diminishes the over-all length (and weight) of the engine. In the 
Thomas-Morse and Sturtevant engines the cylinders are in 
pairs; in the Siddeley "Puma" in threes; in the Hispano-Suiza 
and Bugatti in fours. 

Another variation among airplane-engine cylinders is in the 
form of the cylinder head. In some engines (Hispano-Suiza, 
Napier) the cylinder head is flat and of the same diameter as the 
cylinder barrel. With two equal valves in the head, the maxi- 
mum possible external valve diameter is half the cylinder diame- 
ter minus half the thickness of the bridge between the valves. 

122 



ENGINE DETAILS 123 

It is frequently found that this diameter is insufficient to give 
the desired opening for the admission of the mixture and results 
in a low volumetric efficiency. To remedy this the head may be 
made with sloping sides either without enlargement of diameter 
(Curtiss OX, Mercedes) or with enlargement (Liberty, Lorraine 
Dietrich, Austro-Daimler, Fiat). Another common method of 
meeting this difficulty is by the use of multiple valves. The 
two devices may be used together. 

The most important factor in determining the design of cylin- 
ders is the material employed; the following constructions are in 
use: 

(a) All cast iron, with cylinders either single (Curtiss OX) or in blocks. 
This construction leads to excessive weight. 

(6) Cast-iron barrel, head, and part of jackets, but with aluminum 
sides to the jacket (Bugatti). This is for block construction only. It 
reduces weight and makes the inside of the jacket accessible for machining 
and cleaning. 

(,c) Cast-iron barrel and head, but with sheet metal jackets, either copper 
(Beardmore), steel (Benz) or monel-metal (Curtiss). 

(d) Steel barrels, aluminum heads and jackets. The steel barrel may be 
integral with its head (Hispano-Suiza, Siddeley "Puma") or only a cylin- 
drical shell (Sturtevant). The aluminum jacket may be complete (Hispano- 
Suiza, Sturtevant) or may use the steel barrel as the inner wall (Siddeley 
"Puma"). 

(e) Steel barrel, cast-iron head and steel jacket (Maybach, Benz). 

(J) All steel (Liberty, Packard, Hall-Scott L, Curtiss K, Rolls-Royce, 
Napier, Renault, Lorraine-Dietrich, Fiat, Austro-Daimler, Basse-Serve, 
Maybach). This construction is used more than any other and appears 
to be displacing other constructions. The cylinder is either machined out 
of a solid forging or may be made from drawn steel tubing. The jackets 
are commonly stamped to shape" in two halves along a plane through the 
cylinder axis, and are welded together and to the cylinder. The valve 
ports and guides offer some difficulties as compared with cast cylinder heads. 

Thickness of Cylinder Walls. — If p = maximum gas pressure 
in the cylinder, pounds per square inch, d = cylinder diameter, 
inches, t = wall thickness, inches, and s = allowable tensile 
stress, pounds per square inch, then t = pd/2s, gives the thick- 
ness of metal necessary to withstand the gas pressures. Taking 
s as 5,000 and 14,000 lb. for cast iron and forged steel, the respec- 
tive wall thicknesses for a 5-in. diameter cylinder and for p = 
500 lb. are 0.25 and 0.09 in. With cylinders of small diameter 
the thickness may have to be increased over the calculated 
values to ensure sufficient stiffness and, in the case of cast iron, 



124 



THE AIRPLANE ENGINE 



to offset a possible lack of homogeneity. No allowance need be 
made for wear or reboring as the engines are essentially short- 
lived. 

The clearance space alone of the engine is subjected to the 
maximum explosion pressure; the pressures to which the walls are 
subjected become progressively less from the clearance space to 
the part of the cylinder at the lowest point reached by the top of 
the piston, below which point they become zero. In addition to 
the gas pressures the cylinder walls have to tie the cylinder head 
to the crankcase and shaft bearings and consequently have to 
withstand the maximum gas pressure exerted on the cylinder 
head. The thickness of wall required for this is given by 

pd _ t 

4s ~ 2 



h = jd 2 -p/Tids 



The lowest part of the cylinder consequently does not have to be 
more than one-half the thickness of the upper end. In many 
designs the cylinders are turned of diminishing thickness from 
head to crank end; in other designs (Mercedes, Lorraine Renault, 
Liberty, Curtiss K) the upper part of the cylinder is reinforced by 
stiffening ribs while the lower part is without such stiffening. 
The following table gives some cylinder thicknesses and calcu- 
lated stresses; a maximum gas pressure of 500 lb. per square inch 
is assumed. 



Engine 



Diam- 
eter, 
inches 



Material 



Cylinder- 
head 
thickness 
inches 



Cylinder- 
barrel 

thickness, 
inches 



Calculated 

maximum 

stress, 

pounds 

per square 

inch 



Benz, 230 

Liberty 12 

Curtiss, K-12 

Hall-Scott A5a 

Austro-Daimler, 200 

Renault, 400 

Basse-Serve 

May bach, 300 



5.71 
5.00 
4.50 
5.25 

5.31 
4.92 
6.10 
6.50 



Cast iron 
Steel 
Steel 

Semi-steel 
casting 
Steel 
Steel 
Steel 
Steel 



0.2600 
0.1875 



0.1970 

0.2700 
0.3100 



0.216 
0.156 
0.078 
0.125 

0.138 
0.110 
0.118 
0.110 



6,600 

8,000 

14,400 

10,500 

9,620 
11,200 
12,900 
14,800 



The thickness of the cylinder head is determined mainly by 
considerations of stiffness. It is essential that the valve seats, 
which are located in the cylinder head, should be free from 



ENGINE DETAILS 



125 



deformation and this cannot be secured unless the heads are 
stiff. In cases where the integral cylinder head is backed by an 
aluminum casting (Hispano-Suiza, Napier) the thickness cannot 
be reduced because the heat transfer through the double thickness 
of metal is poor and the exhaust valve seat, which is the hottest 
part of the cylinder, must have sufficient thickness of metal 
around it to conduct the heat away and prevent warping from 
overheating and unequal expansion. The head thicknesses of a 
few cylinders are given in the table above. Steel cylinders are 
generally machined out of solid sheet forgings, but in the case 
of the Liberty engine the cylinders have been made from steel 
tubing ]/2 in. thick, whereby a great reduction is obtained in the 
amount of metal to be removed by machining. 

The valve seats are integral with the heads in steel cylinders 
but have to be cast, or otherwise fastened, into aluminum heads, 
With integral or non-detachable valve seats it is impossible to 
take out the valves without taking off the cylinder or taking out 
the piston. Detachable valve cages have been used, but these 
will always result in imperfect cooling of the valve seat. A com- 
promise adopted in the Beardmore and Austro-Daimler engines is 
to have a detachable valve cage (Fig. 85) 
for the inlet valve, which requires little 
cooling; the exhaust valve can then be 
dropped into the cylinder and withdrawn 
if necessary through the inlet valve seat 
opening. 

The valve ports and valve stem guides 
in all steel constructions are integral for 
each valve and are welded to the cylinder 
head. The guides are provided with bush- 
ings of steel or bronze. The valves must 
work very freely in these bushings and at 
the same time there must be no leakage 
of air or exhaust gas between valve stem and guide; the guides are 
made quite long. 

Many devices have been used for attaching the cylinder to the 
crankcase. The best method is to tie the cylinder, by through 
bolts, directly to the main bearing caps on the crankshaft on 
each side of the cylinder. With two long bolts on each side 
of the cylinder tying the cylinder flange to the main bearing caps, 
the gas pressure on the cylinder head is supported entirely 




Fig. 85. — Detachable 
valve cage (Austro-Daim- 
ler engine). 



126 



THE AIRPLANE ENGINE 



by stresses set up in the bolts and is not transmitted to the crank- 
case, which can consequently be made lighter. As the bolts are 
parallel to the cylinder axis and are symmetrically disposed 
around it, this method of attachment avoids all distortion. One 
pair of bolts is commonly made to serve for two adjacent 
cylinders by the use of dogs through which the bolt goes and 

which are supported equally on 
the flanges of both cylinders 
(Fig. 80). In the Bugatti 
engine (Fig. 67) the through 
bolts are supplemented by studs 
in the crankcase. 

It is only in engines with a 
single row of cylinders that the 
above method of attachment can 
be employed. In Vee and W 
engines other methods must be 
used; the most common is by 
studs in the crankcase which 
pass through the flange at the 
lower end of the cylinder (see 
Fig. 47). In the Curtiss K en- 
gines the steel cylinder is kept 
in place by the aluminum cyl- 
inder head, which is bolted to an 
aluminum jacket cast integral 
with the upper half of the crank- 
case (Figs. 87 and 56). 

Typical Cylinder Designs. — 

Single cast-iron cylinders are sel- 

86.— Cylinder of Hispano-Suiza dom used; a typical example is 

in the Benz 230 engine of Fig. 
77. The barrel and head are cast in one piece. Jackets are of 
die-pressed steel welded on and extending well down towards 
the base flange. They are provided with annular corrugations 
to take care of expansion. Plates welded in position in the jacket 
space above the crown of each cylinder deflect the water to the 
exhaust valve pockets. The cylinder barrels extend 0.39 in. 
below the base flanges into the crank chamber and are held down 
each by four bolts and four dogs. The cylinder walls taper 
from 6.5 mm. at the top to 5.5 mm. at the base. 




Fig. 



ENGINE DETAILS 



127 



An example of the block cast-iron construction is shown in 
Figs. 66 and 67 (Bugatti). The cylinders are in blocks of four. 

Composite steel and aluminum constructions are made up in 
several ways. In the Hispano-Suiza engine (Figs. 86 and 51) 
the steel liner with integral head is threaded throughout the 
entire length of contact with the aluminum. The aluminum 
jacket is a block construction for four cylinders and completely 
surrounds the steel liners so that valve ports go through both the 
aluminum and the steel. For proper cooling it is necessary to 




_J 



Fig. 87. — Cylinder of Curtiss K engine. 



have perfect contact of the two metals at the cylinder head as 
well as the barrel, otherwise warping of the steel head will occur 
and the valve seats will distort. There is no actual contact 
anywhere of water with the steel cylinder. The steel liners are 
attached to the crankcase by bolts. 

A different steel-aluminum block construction is employed in 
the Curtiss K engines. In this case the steel liners with integral 
heads are turned with stiffening flanges (Figs. 87 and 56) on the 
outside, with a packing retaining flange at the bottom and a 
central stud at the top. The upper end of the liner is of slightly 
enlarged diameter and is threaded on the outside. The alumi- 



128 



THE AIRPLANE ENGINE 



num cylinder head is a block casting into which the six cylinders 
are screwed. To ensure intimate contact the threaded stud at 
the center of the liner head, which passes through the head casting, 
is drawn up by a nut. The head casting matches up with and is 
bolted to a flange on the upper end of the aluminum cylinder 
block, which is integral with the upper half of the crankcase. 
Jacket water is in contact with the liner below the combustion 
space only. Water tightness is obtained by packing between the 
lower liner flange and the crankcase. 

Still another construction is used in the Napier "Lion" engine 
(Fig. 73). This engine employs liners with integral heads, 
which are fastened to a four-cylinder aluminum-block head by 
four valve seats (inlet of bronze, exhaust of steel) in each cylinder. 

These seats are screwed into the 
cylinder head. The use of remov- 
able seats is in general objectionable 
as it means less perfect cooling of the 
seats than is possible with an integral 
construction. The barrel jackets in 
the latest design are separate and of 
pressed steel, made in two halves 
welded together and to flanges on the 
cylinders at the top and bottom of 
the water spaces. In the design 
shown in Fig. 73 a common steel 
water jacket is bolted to the head 
casting and makes a joint with each 
cylinder at its lower end by means of a 
rubber ring pressed against a flange on 
the liner by a large circular split nut. 

A steel-aluminum block construc- 
tion in which the liner has no head is 
employed in the Siddeley "Puma" 
engine (Fig. 8). In this case the 
cast in sets of three and the steel liners 
Aluminum water jackets 
The lower 




Fig. 



— Cylinder of Sturte- 
vant engine. 



aluminum heads are 

are screwed and shrunk into them. 

are also cast in threes and are bolted to the heads. 

joint between the aluminum jacket and the liner is made by a 

screwed gland, which squeezes a rubber ring against a shoulder 

on the outside of the liner. The valve seats are of phosphor 

bronze expanded into the aluminum head. 



ENGINE DETAILS 



129 





Fig. 89. — Cylinder of Maybach engine. 





Fig. 90. — Cylinder of Benz 300 engine. 



130 



THE AIRPLANE ENGINE 



Another construction in which a headless steel liner is em- 
ployed is the Sturtevant engine, Fig. 88. The aluminum cylin- 
ders are cast in pairs and are provided with closely fitting steel 
liners; the heads are also of aluminum with inserted valve seats 
and are in pairs. A peculiarity of this construction is that 
the aluminum cylinders are bolted direct to the crankcase; the 
steel liner does not transmit any longitudinal forces and is 
perfectly free to expand. 

A composite steel and cast-iron con- 
struction is used in the Maybach engine 
(Figs. 89 and 81). The steel barrel is 
screwed with a buttress thread into a 
cast-iron head and comes up against a 
soft brass washer which prevents water 
leakage. The water jacket is a machined 
forging and screws on to the cylinder head 
(see details, Fig. 89) where it is sweated in 
position with soft solder; the lower joint is 
kept watertight by a gland and rubber 
ring. The valve-stem guides have cast- 
iron bushings. 

In the Benz 300 engine (Fig. 90) the 
cylinder head is of cast iron and the rest 
of the cylinder is of steel. There are ports 
for two inlet valves and one exhaust valve. 
The steel liner screws into the cast-iron 
head and makes a watertight joint by bed- 
ding into cement in the small groove into 
which the top of the liner goes. 
The all-steel construction is exemplified in the Liberty cylinder 
(Fig. 91). This design has a bumped head and obliquely set 
valves. The jackets and valve ports are welded to the sheet 
cylinder. Details of this welding are shown for the very similar 
Packard engine, Fig. 92. 

In the Austro-Daimler engine (Fig. 93) the cylinders are 
all steel with pressed steel jackets and twin inlet and exhaust 
valves in the cylinder heads. The valve pockets are welded in 
position with the exception of one inlet valve (Fig. 85), which is 
detachable with its seat and guide. The cylinders taper from 
4 mm. at the top to 3 mm. at the middle and increase again to 
4 mm. at the bottom; the jackets are 1 mm. thick. The bottom 




Fig. 91. — Cylinder of 
Liberty engine. 



ENGINE DETAILS 



131 




Fig. 92. — Cylinder head of Packard engine showing welding, W, of the water 

jacket. 




Fig. 93. — Cylinder of Austro- Daimler engine. 



132 



THE AIRPLANE ENGINE 



of each jacket is flanged over and welded to a bevelled flange 
machined on the barrel. 

Pistons. — One of the most important steps in the improve- 
ment of the airplane engine has been the general substitution of 
an aluminum alloy for the cast iron that, until very recently, was 
universally employed for the piston material. The piston must 
be as light as possible in order to keep down inertia forces and at 
the same time must be thick enough in the crown to conduct the 
heat away rapidly from the center to the circumference, where it is 
taken up by the cylinder walls. With cast-iron pistons, reduc- 
tion in weight has resulted in many troubles and especially in the 
burning and cracking of the piston head. Measurements by 
Hopkinson in a Siddeley engine show the cast-iron piston head 
has a temperature of over 900°F. This is borne out by similar 
measurements on the pistons of air-cooled cylinders by Gibson. 





Fig. 94. — Aluminum piston. 



Fig. 95. — Cast-iron piston. 



With aluminum pistons this temperature is reduced to about 
400°F. and at the same time the piston is considerably lighter. 
The two pistons of Figs. 94 and 95 for a 100 by 140 mm. air- 
cooled engine have weights of 1.26 lb. and 1.77 lb., or a reduction 
in weight of 29 per cent by the substitution of aluminum alloy 
for cast iron; these pistons show the temperature difference 
noted above. 

The lower temperature of the aluminum piston has other 
important results. It reduces the rise in temperature of the 
incoming charge during the suction stroke and thereby increases 
the volumetric efficiency of the engine, and it permits the use of a 
higher compression ratio without danger of preignition and 
thereby increases both the capacity and efficiency of the engine. 
The horse power of an engine can be increased at least 5 per cent 
by the use of aluminum pistons. 

The piston should be designed with heat dissipation in mind 
as much as the other piston functions. The heat received at the 
center of the head must pass out radially to the circumference, and 



ENGINE DETAILS 



133 



this should be provided for either by adequate thickness of metal 
from center to periphery or by the provision of radial ribs which 
serve the double function of heat carriers and stiffening members. 
Furthermore, the thickness of the cylindrical wall must not be 
cut down behind the first ring as the heat has to flow downward 
from the crown. 

The friction between the piston and the cylinder is by far the 
largest item of mechanical loss (see p. 24) and should be kept 
as low as possible. Its high value results apparently from the 
partial carbonization of the lubricant clinging to the walls under 
the action of the high gas temperatures and of the slight but 
unavoidable leakage of burning gases past the piston rings. As a 
result, the viscosity of the lubricant is greatly increased. The 




Fig. 96. 





-Slipper piston. 



extent of the friction depends upon: (1) The pressure of the 
piston against the cylinder walls, which governs the thickness of 
the film; the friction appears to be proportional to the average 
loading; (2) the area of the bearing surface; (3) the quantity of 
the lubricant on the walls; the friction increases with this quan- 
tity; (4) the temperature of the walls, which controls the viscosity 
of the lubricant. Of the means adopted to reduce this friction 
loss the most prominent is the cutting away of the piston skirt on 
the sides which do not support side thrust. 

This practice has the further advantage of reducing the weight 
of the piston. An example of such a piston is shown in Fig. 96. 
It will be seen that the piston-pin bosses in this design are 
supported by vertical transverse ribs, which pass within a distance 
from the center of the head of a little more than half of the radius 
of the piston. This is an important feature in keeping the piston 
head cool; the heat absorbed by the central portion of the head 



134 



THE AIRPLANE ENGINE 



can pass down these ribs to the gudgeon-pin bosses and to the 
skirt. This in turn permits the use of a thinner crown, especially 
as the stiffness of the crown is greatly increased by the support 
of the ribs. The thickness of the lubricant film is diminished 
by providing holes in the slippers through which excess oil is 
squeezed out. If both slippers are designed for the same inten- 
sity of pressure, then areas of the two slippers will be different. 
Such a design is shown in Fig. 97; the supporting ribs are no 
longer parallel. 





Fig. 97. — Piston with unequal slippers. 



One of the important troubles with pistons is the slap which 
occurs when the side thrust is transferred from one side to the 
other. The amount of this slap depends on the clearance 
between the piston and cylinder. The cold clearance must be 
larger with aluminum than with cast iron as the expansion is 
much greater. The pistons of Figs. 94 and 95 have cold clear- 
ances of 0.026 and 0.020 in. respectively; the hot clearances for 
both are 0.008 in. The clearances should be greatest at the top 
and where the temperatures are highest and should diminish as 
the bottom of the skirt is approached. With aluminum pistons a 



ENGINE DETAILS 



135 



cold clearance over the top lands of about 0.005 in. per inch 
diameter is necessary. Such a piston will be noisy when cold. 
For cast iron the cold clearance should be 0.003 in. per inch 
diameter at the top, and 0.00075 at the base. The clearances for 
special engines are given in Table 4. 

In order to prevent piston slap, or the opposite danger of 
seizing when hot, the practice has arisen of insulating the piston 
skirt from the ring-carrying portion of the piston. This is most 
readily accomplished by the use of a piston with piston-pin bosses 
carried by ribs (Fig. 97) and with the 
skirt or slippers separate from the 
upper portion of the piston. A design 
of this character is shown in Fig. 98 ; in 
this case a complete skirt is used. Such 
constructions are satisfactory only for 
cylinder diameters up to about 5 in.; 
for larger sizes the piston cooling will be 
inadequate. The clearance necessary 
for the skirts of aluminum divided slip- 
per pistons is about the same as that re- 
quired for the normal cast-iron piston. 

Another method of reducing piston slap is by offsetting the 
wristpin by a small amount, usually not more than J-i in. The 
object of this construction is to cause the piston to tilt slightly 
about the piston pin and therefore to pass progressively instead 
of abruptly from one cylinder wall to the other. Such offset 
is shown in Figs. 96 and 97. 

The composition of the aluminum alloy used in German pistons 
is given below. These pistons are usually die castings and have a 
tensile strength of 28,000 to 31,000 lb. per square inch and 
extension of 4J£ per cent as against about one-half those quan- 
tities for sand castings. 




Fig. 98. — Divided-skirt 
piston. 

















a 

00 


S 
B 


s 




S3 

a 


o 




G 
O 




*« 

JA 


d 

03 

to 

a 


CO 
« 

C 
b0 


c 
S 




o 
O 


a 


o 


w. 


a 


% 


3 


si 


3 
< 


Benz 230 h.p 


6.02 12.13 1.42! 0.31 





Tr. 


Tr. 


80.12 


Austro-Daimler 200 h.p 


7.67 1.33 1.32 0.52 2.21 





Tr 


0.29 


86.66 


Basa6-Selve 270 h.p 


1.90!l5 62 


1.06 


0.45 











80.Q7 













136 



THE AIRPLANE ENGINE 



Piston weights (including piston rings and gudgeon pin) 
vary from 0.19 lb. (Austro-Daimler) to 0.25 lb. (Liberty) per 
square inch of piston area in aluminum construction; for cast 
iron the weight may exceed 0.42 lb. (Maybach). 

Typical Pistons. — Figure 99 is the Maybach cast-iron piston. 
Figure 100 is the Benz cast-iron piston with a thin crown and with 





Fig. 



-Maybach cast-iron piston. 



a hollow conical steel pillar riveted to the piston head and resting 
directly on the middle of the piston pin. The small end* of the 
connecting rod is cut away to avoid interference with this pillar. 
With this arrangement the gas pressure is transmitted in the line 
of the connecting rod and there is no bending moment on the 
wristpin. The ribbed aluminum construction is shown in Fig. 



r ~ n 






*m 














f 




Fig. 100. — Benz 230 cast-iron piston. 

101 for the same engine. The top clearances for the two pistons 
are 0.02 and 0.03 in. respectively; the bottom clearances are 
0.004 and 0.014 in. respectively. The weights are 6.72 lb. and 
4.90 lb. respectively, complete with rings and setscrews but 
without piston pins; the saving in weight is 27 per cent. The 
ribless construction is typified by the Liberty engine piston shown 
in Fig. 102; oil grooves are provided on the piston skirt. 



ENGINE DETAILS 



137 



Piston rings are of dense gray cast iron, fully machine-finished, 
peened on the inner curved surface and exactly ground to size 
upon the outer curved surface. With very narrow rings semi- 
steel is used. The rings may be either of the concentric or 
eccentric types. The ends are commonly chamfered at an angle 
of 30 to 40 deg. but stepped ends are also used; the gap when in 
the cylinder is about ^40 the diameter of the piston. Three 





^ 


^~\ 




(hi 




N 






wk 


&Mi 




■ \ 

2Sr & 





Fig. 101. — Benz 230 aluminum ribbed piston. 



rings are commonly used, but four rings are found occasionally. 
Two narrow rings are sometimes used in one groove. A scraper 
ring near the bottom of the skirt is sometimes used to clear excess 
oil from the cylinder walls; the same result is obtained by the 
use of perforations through the skirt. In some pistons (Figs. 
99 and 100) the lowest ring acts as a scraper and has a groove 
below it through which small holes are drilled to the interior of the 
piston to drain away any excess of oil. 



138 



THE AIRPLANE ENGINE 



Piston or Gudgeon Pin. — The piston pin is usually of steel,, 
machined to size, case-hardened and ground. It is always 
hollow. It is most commonly fully floating, that is, it has bearing 
in the end of the connecting rod as well as in the piston bosses, 
with some end motion as well.. The pin will then rotate and 

local wear will be avoided. The 
bushing in the connecting rod 
is also floating in many engines. 
On the cold motor the pin 
should be a mild driving fit in 
the bosses and a running fit in 
the connecting-rod bushing. 
When warm the aluminum 
bosses expand more than the 
bushing and the pin becomes 
free. With standard piston 
types, as in Figs. 99-102, the 
piston pin is comparatively long 
and is subjected to considerable 
bending stress. By the adoption 
of the slipper piston (Figs. 96 
and 97) or the divided skirt 
(Fig. 98) the bosses are brought 
closer together and the pin is 
shortened. It can consequently 
be made of smaller diameter and 
lighter and if fully floating~will 
show no wear. 

Connecting Rods. — In vertical 
engines the connecting rods are of 
uniform (non-tapering) circular 
or I-section, with solid small 
ends and marine type big ends. An example of the circular 
section (Benz 230) is shown in Fig. 103; the I-section, assembled 
with the piston (Siddeley "Puma"), in Fig. 104. The small end is 
usually provided with a bronze bushing, although in the Maybach 
engine this is replaced by a perforated cast-iron floating shell 
0.124 in. thick. The large end has a babbitted-bronze shell. 
The tubular rods used in several German engines are sometimes 
provided (Fig. 103) with a centered internal pipe for lubricating 
the small end. The Benz rod has a number of radial holes 




Fig. 102. 



-Liberty ribless aluminum 
piston. 



ENGINE DETAILS 



139 



drilled in the big end to reduce weight and has the top of the 
small end cut away to permit direct application of the gas 
pressure load on the crankpin through a pillared piston as in 
Fig. 100. 

In Vee engines several connecting rod arrangements are used. 
In the Curtiss OX and V, Sturtevant and Thomas-Morse engines 






Fig. 103. — Benz tubular connecting rod. 

the cylinders in the two rows of the Vee are staggered so that the 
connecting rods do not lie in the same plane. With this arrange- 
ment the big ends of each pair of cylinders lie side by side on the 
same crankpin. Such an arrangement results in an increase in 
the over-all length of the engine. 



140 



THE AIRPLANE ENGINE 



With the cylinders of each pair opposite one another, as in 
the general practice with Vee engines, the connecting rods are in 
the same plane and special arrangements must be made to connect 
them both to the crankpin. Two arrangements are in use, the 
forked rod and the articulated rod. The forked rod is used in the 
Liberty, Hispano-Suiza, Packard, and Fiat engines. Each pair 
of rods consists of a plain rod and a forked rod. The forked rod 
clamps the big end bronze bushing; the plain rod works on the 





/re^fesm 




Fig. 104. — Siddeley "Puma" I-section connecting rod. 

outside of this bronze bushing between the two forks of the forked 
rod. The rods for the Liberty engine are shown in Fig. 105. The 
bearing is prevented from rotating in the forked rods by dowel 
pins. In the Fiat engine the bottom ends of the fork are fastened 
together. 

In the articulated rod assembly a master rod is used and a 
short rod is attached to a pin which is held in the upper half of 
the big end of the master rod. Ordinarily the master rods are all 
placed on one side of the Vee, but occasionally (Renault) the 
master rods alternate with short rods. A good example of a 



ENGINE DETAILS 



141 





/a\ 



Fig. 105. — Forked connecting rods of Liberty engine. 




Fig. 106. — Master rod of Benz 300 engine. 



142 



THE AIRPLANE ENGINE 



master rod is shown in Fig. 106 for the Benz 300, 60-deg. Vee 

engine. The rod is tubular; the pin for the small rod is held by 

two clamp screws. In the Renault engine (Fig. 107) the pin for 
the small rod is of the same diameter as the 

LJI 1T\ piston pin so that both ends of the small rod 

are alike. 

In W engines the articulated rod is most 
common. The Napier "Lion" uses a central 
master rod, on each side of which is mounted 
an articulated rod (Fig. 108) carried on pins 
fixed in lugs integral with the big end of the 
master rod. The main rod is of I-section; the 
side rods are tubular and carry bronze bush- 
ings at both ends. Each side-rod pin is 
tapered at one end, fits into a tapered hole in 

the corresponding lug, and is drawn up tight by a bolt screwing 

into the pin at the taper end. 

In the Lorraine W engines the two outer rods bear on the 

cylindrical outer surface of the big end of the master rod, the 




Fig. 107. — Articu- 
lated connecting rods 
of Renault engine. 




Fig. 108. — Articulated connecting rods of Napier "Lion 



bearing slippers covering less than half the circumference. Two 
circular !' steel' rings hold the two halves of the big end of the 
master rod together and hold the slippers of the outer rods to the 
outer surface of the master rod. 




ENGINE DETAILS 143 

The articulated arrangement of connecting rods suffers from 
some disadvantages as compared with an arrangement in which 
the connecting rods are always radial to the crankpin. The 
short rod is materially shorter than the master rod — usually at 
least 20 per cent shorter — and consequently causes greater 
angularity of that rod and increased side thrust in the cylinder. 
The explosion pressures transmitted along the short rod do not 
act directly on the crankpin but impose stresses on the master 
rod which under unfavorable conditions may be serious. For 
example, with a 90-deg. Vee 
and with explosion pressure 
reached 30 deg. before dead 
center in the short-rod cylinder, 
the force acting on the master 
rod CD (Fig. 109) would be 
directed along the line AE. 
The reaction of this force at 
C can be readily found and, 
treating the master rod as a Fig. 109. — Diagram of articulated con- 
cantilever loaded with this re- necting rods. 
action force at C and held on the crankpin, the stress at any 
section of the rod can be ascertained. 

Connecting rods up to the present time have always been 
made of steel. The use of forged aluminum rods would mate- 
rially reduce the weight of the reciprocating parts and the bearing 
pressure at the big end. 

Crankshafts. — Crankshafts are made of alloy steel (nickel 
or chrome-vanadium) and are usually forged in one piece. An 
exception to this is the Bugatti eight-throw shaft which is made 
in two lengths. In airplane practice there are usually bearings on 
both sides of each throw; this gives great stiffness to a light shaft. 
The arrangement common in automobile engines of two or more 
throws between main bearings has been employed by the Sturte- 
vant, Thomas-Morse and Duesenberg engines and is still used 
in the Curtiss K engines (see Fig. 55). The long crankarms 
(between the first and second, and between the fifth and sixth 
cranks) of this engine have centers of gravity which do not 
coincide with the axis of the shaft (as in four-cylinder engines) 
and consequently produce an unbalanced moment about the 
crank axis. This can be balanced by a counterbalance weight 
between the two center crankpins which are in line. The addi- 



144 THE AIRPLANE ENGINE 

tion of such a counterbalance weight sets up an undesirable 
bending moment on the long central crankpin and also puts 
more load on the two central main bearings, which have to resist 
the moments created by these unbalanced masses. To eliminate 
these objectionable conditions it is desirable to balance directly 
the masses of the two long crankarms. This is accomplished in 
the Curtiss K engines (Fig. 55) by applying balance weights 
directly to the long crankarms. To obtain the greatest possible 
balancing effect with the least weight, aluminum spacers are 
inserted between the steel balance weights and the crankshaft, 
the balance weights being held to the crankshaft by steel bolts. 

The crankshaft and pins are always made hollow and holes 
are drilled through the crank webs for oil passages connecting the 
hollow crankpins and journals. The open ends of the shaft and 
crankpins are plugged by screw plugs (Hispano-Suiza) or by 
discs or caps which are expanded or brazed into place, or in some 
cases (Liberty, Siddeley "Puma") are held in position by bolts 
which tie together a pair of caps. In this last case the bolts can 
be used to obtain rotational balance, the method being to use 
special bolts thickened in the middle. The caps in different con- 
structions are of duralumin, gun metal, steel and other metals 
and may be used with or without gaskets. 

Main bearings and crankpin bearings are nearly always of 
babbitted bronze. Occasionally a ball bearing is used at one 
end; at the rear end in the Hispano-Suiza; at the front end in the 
Fiat. In the Napier ' 'Lion " with three cylinders on each crankpin 
and with a heavy big end, the main-bearing pressures are very 
high and roller bearings are used throughout. Double-thrust 
ball bearings are usual and are placed just behind the propeller 
hub. 

The pressure on main bearings is high and demands con- 
siderable oil circulation, not only for lubrication but also for 
cooling. The babbitt tends to flake off unless the bronze has been 
tinned before casting the babbitt, in which case a perfect bond can 
be obtained; mechanical holding of the babbitt by holes, dove- 
tails or screw threads is generally found unsatisfactory. 

The bearing pressure in a six-throw, seven-bearing crankshaft 
is greatest at the center main bearings because the crank throws 
on the two sides of it are in line so that the dynamic loads imposed 
on the two cranks are in phase. Consequently the center main 
bearing is often made longer than the intermediate bearings to 



ENGINE DETAILS 145 

dimmish the intensity of pressure. In Rolls-Royce and Fiat 
engines the center bearings are 60 per cent longer than the 
intermediate bearings. In the Liberty engine the maximum 
load on the center main bearing is 7,700 lb. or 1,675 lb. per 
square inch of projected area; the mean unit bearing pressure is 
1,265 lb. per square inch. As the rubbing velocity is 19.5 ft. 
per second, the friction work is F = f X 19.5 X 1,265 ft.-lb. 
per second, where / is the coefficient of friction. On the inter- 
mediate main bearings of this engine the maximum load figures 
out as 7,250 lb. or 1,580 lb. per square inch of projected area; the 
mean unit bearing pressure is 700 lb. per square inch. The end 
main bearings receive loading on one side only and show a maxi- 
mum load of 4,025 lb. or a maximum unit pressure of 815 lb. 
per square inch and a mean unit pressure of 610 lb. per square 
inch. 

The crankpin bearing pressure for the Liberty engine has a 
maximum value of 4,980 lb. or 932 lb. per square inch of projected 
area; the mean unit bearing pressure is 642 lb. per square inch. 
The crankpin pressures used in the German engines are somewhat 
lower, ranging from a mean unit bearing pressure of 402 lb. per 
square inch in the Austro-Daimler to 585 lb. per square inch 
in the Maybach. The crankpin pressure is mainly due to 
inertia and centrifugal forces — the gas pressures have com- 
paratively little effect. This is evidenced by the fact that the 
wear on crankpins bearings is on the side remote from the piston, 
that is, on the side subjected only to inertia and centrifugal forces. 

Crankshafts are subjected to stresses which vary rapidly 
in sign and magnitude and consequently are especially liable to 
fail from fatigue of material. The weakest point is generally at 
some place where there is a sudden change in cross-section and 
poor distribution of stress. It is particularly important that the 
fillets at the junctions of the crankpins and journals with 
the crank webs should be of adequate size. Tests to determine the 
desirable size of the fillet have been conducted recently in 
England; they show that the steel is materially weakened if the 
fillet is less than % in. radius. 

In the discussion of torque on page 47 it has been shown that 
the maximum torque at the propeller end of the crankshaft of a 
six-cylinder engine is actually less than the maximum value at 
the rear crankpin. Consequently there is no need for any 
increase in diameter of the crankshaft from rear to front in that 

10 



146 THE AIRPLANE ENGINE 

case. The free end of the crankshaft is subjected to much more 
severe conditions than the propeller end. The free end is, as it 
were, wound up when the maximum torque is applied to it and 
released when the torque diminishes. At certain speeds this 
alternate winding up and release may coincide with a natural 
period of vibration of the crankshaft, and in that case the shaft 
will vibrate excessively and the reciprocating masses attached 
to it will also vibrate and impart their vibration to the whole 
structure. Such torsional vibration could be reduced by the use 
of a flywheel on the free end. It has given much trouble in 
various six-cylinder engines and has been largely responsible for 
the failure of engines with eight crank throws. 

With six-cylinder engines of 200 to 300 h.p. the freely vibrating 
shaft has a frequency which is usually about 6,000 vibrations per 
minute; for four-cylinder engines this frequency is higher, and in 
single-crank radial or rotary engines it may be as high as 20,000. 
The period of vibration can be determined by striking a series 
of light blows at regular intervals; the vibrations will increase 
markedly when the frequency of the blows coincides with the 
natural period of the shaft. In a six-cylinder engine the impulses 
are three per revolution so that the dangerous speed for a shaft 
with vibration frequency is 6,000 per minute of 6,000 -5- 3 «= 
2,000 r.p.m. The next most dangerous speed would be 1,000 
r.p.m. With eight-cylinder engines the dangerous speeds would 
be Y^ and }i the vibration frequency. 

The addition of counterweights to the crankshaft as a means 
of obtaining rotary balance is of value mainly in reducing bearing 
pressures. The centrifugal force arising from unbalanced 
rotating weights acts radially from the center and produces con- 
siderable bearing pressures. Under airplane-engine conditions a 
lower total weight is obtained by omitting counterbalances and 
giving the bearing sufficient area and stiffness to support the 
centrifugal forces. With higher speeds of rotation the need for 
counterbalance weights increases. 

Crankshafts are drop forgings and are usually made in dies 
when the quantity warrants it; in other cases they are cut from 
large billets. The dies may be made of cast iron when the 
number of forgings required is small; for quantity production 
they are of steel. 

Strength of Shafts. — Shafts should be designed for their 
strength in shear. For a solid circular shaft of diameter d in. 






ENGINE DETAILS 



147 



subjected to a bending moment M in pound-inches and a torsion 
T, also in pound-inches, and with a maximum permissible intensity 
of shearing stress at the outer surface of the shaft of / lb. per 
square inch, 

\/M 2 + T 2 



5.1 



/ 



For a hollow shaft of outside diameter d 2 and inside diameter di, 
the equation becomes 

d 2 * 



d* 



dS K 1 VM 2 + T 2 
- = 5.1 j- 




Fig. 110. — Propeller hub of Hispano-Suiza engine. 



With M = 24,500, T = 54,000, and/ = 16,000, these equations 
give d = 2% in. and if d% is assumed to be 3 in. d\ is 2.275 in. 
Under these conditions the hollow shaft will weigh 56 per cent 
as much as the solid shaft. A hollow shaft of still larger outside 
diameter would be lighter and stiffer but would require larger 
and heavier bearings and would result in increased rubbing 
velocities at the bearings and increased friction. 

Propeller hubs in American practice are mounted on a tapered 
extension of the crankshaft. The Hispano-Suiza hub (Fig. 110) 
is a good example of standard practice. It is keyed to the 
engine shaft, which is given a taper of 1 in 10, and is threaded at 
the end to receive a long nut which is used for forcing the hub 
on the taper. The inner flange is integral with the tapered hub; 
the outer flange has splines which fit in grooves on the outer end 



148 



THE AIRPLANE ENGINE 



of the hub and permit axial movement of about 1 in. for adjust- 
ment to the thickness of the propeller, which is held between the 
two flanges. Eight bolts hold the flange and the propeller 
together. Rotation of the hub on the engine shaft is prevented 
by a key. The long nut is held in position by a locknut or a 
locking pin. 

The Benz engine (Fig. Ill) employs a hub which is bolted to a 
flange at the end of the crankshaft and has an outer flange which 
fits on the splined end of the hub. 

Crankcases. — The crankcase has to serve several functions: 
it has to tie various parts of the engine together; it has to with- 
stand stresses due to gas pressures, and bending moments due to 
unbalanced forces; it contains the lubricating oil; it supports 




Fig. 111. — Propeller hub of Benz engine. 



various auxiliaries; and it has to support the engine as a whole. 
The stresses in the crankcase are chiefly of the two kinds sug- 
gested above. The explosion pressure, acting on the cylinder 
head, puts the cylinder under tension and this should be supported 
as directly as possible by connecting members from the cylin- 
der to the corresponding lower main crankshaft bearings. In 
vertical engines this can be accomplished very satisfactorily by 
the use of through bolts from the lower cylinder flange to the 
lower bearings (see Fig. 112), using transverse webs in the 
upper crankcase as distance pieces. In Vee and W engines 
this construction is not possible and it is necessary to fasten the 
cylinders and the lower bearings to the upper crankcase, and 
transmit the tension through the transverse webs to the lower 
bearings. The lower half of the crankcase is sometimes cast as a 
unit with the lower bearings but this practice has nothing to 
recommend it. The assembly of cylinders and upper crankcase 
should sustain all the stresses due to gas pressures. 



ENGINE DETAILS 



149 






The unbalanced centrifugal and inertia forces acting through 
the main bearings subject the crankcase to bending moments 
which change continuously in direction and magnitude. It is 
necessary that the crankcase should have sufficient stiffness to 
withstand these bending moments without objectionable deflec- 
tions. For this purpose a webbed box structure has been found 
most satisfactory. It is easily possible to obtain the necessary 
stiffness by utilizing the upper crankcase only as a stressed 
member. The lower crankcase is preferably used only as an oil 




Fig. 112. — Transverse section of crankcase. 

container and a support for oil pumps and other auxiliaries. 
This practice can be seen in the Hall-Scott L (Fig. 64), Bugatti 
(Fig. 67), Curtiss K and OX (Figs. 55 and 59), Napier "Lion" 
(Fig. 72), Maybach (Fig. 81), and Benz engines (Fig. 78). In 
other engines the lower crankcase carries also the lower halves of 
the end main bearings, as in the Hispano-Suiza (Fig. 51) and Sid- 
deley "Puma" engines (Fig. 9). These arrangements permit of 
easy accessibility. In the Liberty engine (Fig. 47) the lower 
crankcase has transverse webs and the crank chamber is divided 
into six separate chambers; a similar construction with double 
transverse webs is employed in the Fiat engine (Fig. 76). 



150 THE AIRPLANE ENGINE 

The transverse webs are often cut away in places for lightness. 
Aluminum alloy is universally used for crankcases. 

The lower crankcase serves as an oil sump. The earlier 
practice of keeping a considerable body of oil (wet sump) in the 
bottom of the crankcase is now being superseded by the dry sump 
which is kept drained by a scavenger pump or pumps. Wet sumps 
are shown in the Benz engine (Fig. 78), Hispano-Suiza (Fig. 51), 
Hall-Scott L (Fig. 64), and Curtiss OX engines (Fig. 59). In 
the last case the sump is separated by drainage plates from the 
rest of the crankcase. The advantage of the dry sump is in 
avoiding drowning the cylinder with oil in case the engine 
operates momentarily upside down or in any posture approximat- 
ing that position. The drainage point of the sump is usually 
in the middle, but in some cases the scavenger pumps take oil 
from both ends (Liberty, Napier "Lion")- Oil cooling is carried 
out in the lower crankcase in the Austro-Daimler engine by 
casting outside cooling ribs running longitudinally along the 
bottom of the crankcase and attaching a sheet of aluminum in 
such a way as to form an air duct along the whole underside of the 
engine. In the Basse-Selve engine an oil cooler with air tubes is 
fastened to the bottom of the crankcase but has no direct com- 
munication with the inside. In the Curtiss K engine (Fig. 55) 
oil cooling is effected inside the crankcase by the jacket water on 
its way to the pump; this arrangement serves also to heat up the 
oil quickly after starting the engine and puts the lubrication 
system into normal operation earlier than would otherwise be 
possible. 

German airplane engines usually have provision for air cooling 
of the crankcase. In the Benz engine (Fig. 77) the support- 
ing webs for six of the main bearings form air passages trans- 
versely across the engine. Two of these serve as air intake 
passages to the two carburetors, which are thereby supplied with 
heated air. In addition the lower crankcase is traversed by 18 
aluminum tubes, 30 mm. in diameter; air is scooped into the 
tubes through an aluminum louvered cowl on one side of the engine 
and discharged through a reversed cowl on the other side. 



CHAPTER VII 
VALVES AND VALVE GEARS 

Location of Valves. — The diversity of valve locations which 
characterizes automobile practice is not found in modern airplane 
engines. L-head and T-head arrangements have been supplanted 
by overhead valves which permit the simplest form of cylinder 
head and water jacket, shortest and most direct passage of the 
gases, and, with overhead camshafts, a considerable simplifica- 
tion in the valve gear and a reduction in the number of rubbing 
or contact points. The valves may either be seated in a flat 
head in which case the stems are parallel to the cylinder axis or 
the head may be domed, in which case the valve stems are 
inclined to the cylinder axis. 

Valve Lift. — Valves are always of the poppet type with bevelled 
seats. They are opened by cams which operate either directly 
or indirectly; they are closed by springs. The valve (Fig. 113) 
has a face which is usually about 25 per 
cent wider than the seat on which it 
closes and a stem which passes through 
a long guide (often provided with a 
bushing) and which connects with the 
valve by a rounded fillet. The bevel 
of the valve face is usually 45 deg. but 
30 deg. is sometimes used. The width 
of the valve face must be small to 
ensure gas tightness and is usually 
about one-fourth the lift of the valve. 
The width of the valve seat is usually 
less than 0.1 in. The free area for 
the passage of the gas through the 
fully opened valve may be taken approximately as irdh where 
d is the smallest diameter of the bevelled valve face and h 
is the lift. This area should not be greater than the free open- 
ing through the valve seat. Neglecting the area occupied 

by the valve stem, xdh = ,d 2 , or h = ^id, gives the lift which 




Fig. 



113. — Typical airplane 
engine valve. 



151 



152 THE AIRPLANE ENGINE 

makes the gas passage area equal to the free opening through the 
valve; usually h varies from one-fifth to one-sixth of the outside 
valve diameter. Values will be found in Table 4. 

With a valve lift of one-quarter of its diameter the gas flow 
for a given pressure drop is found by experiment to be about 67 
per cent of the flow through an unobstructed port 1 ; with a lift 
of one-half the diameter this is increased to from 80 to 90 per cent. 
These " coefficients of efflux" are found to be the same for all 
pressure drops, and for valves of different sizes at equal lifts 
expressed in per cent of their respective diameters. The experi- 
ments were carried out with continuous flow which presumably 
would give results differing from those actually occurring under 
the operating conditions of intermittent flow. The earlier in- 
vestigations of Lucke 2 indicate coefficients of efflux lower than 
those given above; the variation with lift is probably of the right 
order of magnitude. 

The volume of gas passing the inlet valve per unit of time is 
approximately equal to the piston displacement in that time; the 
volume of gas passing the exhaust valve is from two to three 
times as great. If the mean piston speed during the suction 
stroke is s in. per second and the piston diameter is D in. the 
mean gas velocity V in feet per second past the valve is given by 

^D 2 s = UVirdh 
4 

or 

V — *c 77 ft. per second 

where d and h are in inches. This velocity is approximate, as the 
equation assumes the cylinder to fill with a mixture at atmos- 
pheric pressure and temperature. As the piston speed is varying 
throughout the stroke, the gas velocity will vary and will have a 
maximum value which is nearly twice the mean value. 
For the Liberty engine the mean gas velocity is 189 ft. per second; 
for most engines it varies from about 150 to 200 ft. per second. 
The pressure drop past the valve to obtain this velocity can 
be obtained with sufficient accuracy from the equation V 2 = 2gh, 
where h is the pressure drop measured in feet of air. To convert 
this to a pressure drop, i } measured in inches of water, using the 

1 Lewis and Nutting, 4th Ann. Report, Nat. Adv. Comm. Aeronautics, 1918. 

2 Trans. A. S. M. E., 1905, Vol. 27. 



VALVES AND VALVE GEARS 153 

conversion factor, 1 in. of water = 5.2 lb. per square foot, we have 

h = 5.2 X i X v 

where v is the volume of 1 lb. of the explosive mixture. The 
quantity v is obtained from the perfect gas equation 
v = RT/p = 52 T/p. 

■ At ordinary atmospheric pressure and temperature, v has a value 
of about 13. 

V 2 = 2g X5.2 X 13 Xi 
and 

• = y2 
1 4,350 

For gas velocities from 150 to 200 ft. per second the corresponding 
pressure drops are from 5.2 to 9.2 in. of water. 

The pressure drop through the valve is important as affecting 
the pressure in the cylinder at the end of the suction period and, 
thereby, the volumetric efficiency. This pressure is controlled 
also by the frictional resistances to the flow through carburetor, 
manifold and gas ports, by the time and rate of valve opening 
and closing, and by other factors. At midstroke, when the piston 
velocity is a maximum, the gas velocity past the valve will be a 
maximum; if its value is twice the mean velocity the pressure 
drop will be four times the mean value. In the Liberty engine 
the pressure drop corresponding to a mean velocity of 189 ft. per 
second is 8.2 in. of water; the pressure drop past the valve at 
midstroke would then be about 4 X 8.2 = 33 in. of water. Fric- 
tional resistances will increase this quantity. During the latter 
half of the stroke the quantity of gas passing the valve will be 
greater than in the first half on account of the existence of this 
greater vacuum. That is, during the first half of the stroke a 
vacuum is being created in the cylinder; during the last half this 
vacuum is being filled up. Near the end of the stroke the valve 
is closing which cuts down the area for gas flow and limits the 
filling up process. In most engines (see p. 175) the final closure 
of the valve does not occur till after the dead center and the 
additional time so obtained for the admission of the charge has 
important results in increasing the weight of charge admitted. 
The continued flow of gas into the cylinder past the dead center 
position is due not only to the existence of a vacuum there but 
also to the inertia of the column of gas in the induction system. 
It is not desirable to delay the final closure of the valve long 



154 



THE AIRPLANE ENGINE 



enough to establish atmospheric pressure in the cylinder; this 
procedure would not increase the volumetric efficiency because 
of the diminished volume of the charge resulting from the 
return of the piston. The actual closing of the valve is some- 
times delayed till the piston has returned 10 per cent of its stroke 
so that the volume of gas at the beginning of compression is 
correspondingly reduced and the compression ratio is less than 
the cylinder dimensions would indicate. 

The gas flow area past the valve has been seen to be propor- 
tional to dh, or since h is proportional to d, the area is proportional 
to d 2 . The larger this area the lower will be the gas velocity, 
and the less the frictional resistances and the pressure drop. 




(d) (e) (f) 

Fig. 114. — Valve arrangements. 



With flat-headed cylinders and with mo increase in diameter 
in the combustion space the valve diameter must be considerably 
less than half the cylinder diameter; in the Hispano-Suiza 180 it 
is 42 per cent, in the Hispano-Suiza 300 it is 40 per cent. With 
enlarged heads the ratio of valve diameter to cylinder diameter 
can easily be made 50 per cent as in Liberty engines, or even 
55 per cent as in the Curtiss VX engine. 

To obtain a larger valve area, dual valves may be employed. 
The Benz 200 has dual valves of 1.693 in. diameter as against one 
valve of 2.205 in. diameter in the Hispano-Suiza 300 engine of the 
same cylinder diameter; the possible gain in valve area is 17J^ 
per cent. A comparison is shown in Fig. 114 of the maximum 
valve areas for the Hispano-Suiza 300 engine, a with single 
valves, b with dual intake and dual exhaust valves and c with dual 
intake and single exhaust. The single valves are of the same 



VALVES AND VALVE GEARS 155 

diameter as in the actual engine; the width of bridge between 
valves is kept constant and the clearance from the cylinder walls 
is the same as in the actual engine. The maximum dual valve 
area in b is 19 per cent and in c is 45 per cent greater than the 
single valve area in a. In Fig. 114 d the single inlet valve 
is shown of the same area as the dual exhaust valves; with this 
arrangement both inlet and exhaust valve areas are increased 23 
per cent as compared with a. The increase in valve area obtain- 
able by the use of dual valves becomes greater as the cylinder dia- 
meter increases since it is not necessary to change the width of 
bridge between valves or the cylinder wall clearance with change 
in cylinder diameter. For example, with a cylinder diameter of 
7 in. and bridge width 0.69 in. and Wall clearance 0.206 in. the 
use of dual valves increases the area 24 per cent as compared 
with 19 per cent gain in the Hispano-Suiza 300 with cylinder 
diameter 5.51 in. and the same bridge width and clearance. 

A further development along the same lines is the use of 
triple inlet and exhaust valves as in the 800-h.p. Sunbeam Sikh 
engine. This arrangement leads to a decrease in valve area 
except for large cylinders. With a 7-in. cylinder and the same 
bridge width and clearance as above the decrease in valve area 
with triple valves in line (Fig. 114 e) is 35 per cent as compared 
with dual valves; with the valves arranged in a circle (Fig. 114/) 
the decrease in valve area is 16 per cent. The latter arrange- 
ment would require a more complicated valve gear than is 
necessary with the three valves in line. With larger cylinder 
and smaller bridge width the comparative showing of the triple 
valves would improve. 

The comparisons just presented between multiple and single 
valves are based on the assumption of the same lift expressed in 
per cent of diameter. If dual valves are used and the actual lift 
is kept the same as for the corresponding single valve, the gas 
flow through the multiple valves will be increased by about 20 
per cent over the figures given; for example, the dual valves of 
Fig. 114 b will increase the gas flow about 40 per cent above that 
through the single valve for a given pressure drop. The diam- 
eters of the dual valves can be reduced, without decreasing the 
gas flow below that through a single valve, by maintaining the 
same lift as for a single valve. For example, a 2.5-in. valve with 
25 per cent lift (0.625 in.) gives the same flow as two 1.75-in. 
valves with 25 per cent lift (0.44 in.) or two 1.5-in. valves with 



156 THE AIRPLANE ENGINE 

0.625-in. lift (42 per cent). This last arrangement has certain 
advantages; the area of valve presented to gas pressure is only 
72 per cent of the area of the single valve and the weight of the 
two valves will be only 56 per cent of the weight of the single 
valve, assuming the weights to vary as d 2 - 5 . With valve spring 
tensions proportional to valve weights the power required to 
operate these dual valves will be less than half the power required 
to open the single valve. Other advantages are that a larger 
proportion of the cylinder head can be jacketed because it is not 
occupied by the valves; that the valve cooling will be better 
because the circumference of the two valves is 20 per cent greater 
than that of the single valve and the distance travelled by the 
heat is only 60 per cent as great; and the distortion of the valve 
will be less in consequence both of lower temperature and smaller 
diameter. 

The area of the exhaust valve opening affects the volumetric 
efficiency since it determines the pressure of the residual gases 
in the combustion space at the time of opening of the inlet valve 
and determines the back pressure on the piston during the exhaust 
stroke. The main problem of the exhaust valve is heat dissipa- 
tion. The exhaust valve is heated during the explosion stroke 
and by the exhaust gases as they pass out. The greater the 
velocity of the gases the greater is the heating of the valve. Heat 
abstraction from the valve is principally by conduction to the 
seat and thence to the jacket water, but is also by conduction to 
the stem and thence to the guide. The inlet valve gives no 
trouble from this source as it is cooled by the entering charge. 
Increase of exhaust valve area by increasing its diameter is 
objectionable because it results in increased temperature of the 
valve. The heat received by the valve is approximately pro- 
portional to the square of its diameter while the area of the heat- 
abstracting seat is proportional to the first power of the diameter. 
Furthermore, as valve diameter increases the mean distance the 
heat has to travel increases and this results in increased valve 
temperature and also in valve warping. The maximum practica- 
ble size of exhaust valve seems to be about 3 in. diameter. A 
limit to valve size is also set by the valve weight, which increases 
about as d 2 - 5 while the area increases as d 2 . As the valves are 
closed by springs, the dimensions of the springs have to increase 
rapidly with increase of valve diameter and lift in order to give 
sufficiently quick closure. The desirable method of increasing 



VALVES AND VALVE GEARS 



157 



exhaust valve area is therefore by the use of multiple valves and 
not by increase in diameter of a single valve. The arrangement of 
Fig. 1146 with dual exhaust valves and single inlet valve, is often 
employed (Siddeley "Puma," Benz 300, ABC Dragon-fly), 
while the reverse arrangement of two inlets and one exhaust 
valve is seldom used (Bugatti). Ordinarily, if dual valves are 
used, they are used both for inlet and exhaust. 

In addition to providing the maximum free area through the 
valves it is important that the form of the passage through which 
the gas is flowing should be such as to offer minimum resistance 
and as far as possible to permit smooth flow without the forma- 




Inlet 




Exhaust- 



Fig. 115. 



-Trumpet-shaped 
valve. 



Fig. 116. — Valves of Basse-Selve engine. 



tion of eddies. Especially is this true for the inlet valve as it 
will influence materially the volumetric efficiency of the engine. 
The limitations inherent in an engine as light and compact as an 
airplane engine will generally prevent the adoption of ideal 
forms of passage, but some improvements over common practice 
(as represented in Fig. 113) are possible. One of these is in the 
shape of the valve. If the valve is formed with a trumpet head 
which flows with a large radius into the stem as in the Siddeley 
"Puma" inlet valve (Fig. 115), the gas flow lines will be consider- 
ably improved; such valves should be hollowed out to reduce the 
weight. The approach of the gas to the valve is usually smoother 
when the valves are inclined than when they are vertical; compare, 
for example, the Basse-Selve (Fig. 116) with the Curtiss K (Fig. 



158 



THE AIRPLANE ENGINE 



87), both of which are good examples of smooth passages without 
sudden enlargements. In the Liberty engine (Fig. 91) the 
form of the gas passages is such as to set up eddies. On the 
exhaust side the need for adequate water-jacketing of the valve 
seat and guide will generally lead to a less favorable form of gas 
passage than on the inlet side. This is especially noticeable in the 
Basse-Selve engine (Fig. 116), in which the water jacket is carried 
all round the whole length of the valve stem guide; in most 
engines such complete jacketing is not attempted. 

The cross-section area of inlet pipes and ports should be 
approximately the same as the valve area so as to avoid loss 
of head due to change of cross-section. With an engine using 0.6 



340 



2 
o30 



!20 



















/^ 


, B area 


O.lsq-in. 















1200 1600 2000 

Revolutions per Minute 



2400 



Fig. 117. — Variation of engine 
power with valve lift. 



50 



|40 
o 

Cl 













J 




O^sq-m. 



















1200 , 1600 2000 

Revolutions per Minute 



2400 



Fig. 118. — Variation of engine 
power with valve diameter. 



lb. of fuel per horse-power hour and 15 lb. of air per pound of fuel, 
the volume of charge entering the cylinder will be approximately 
120 cu. ft. and of the exhaust gases 250 to 300 cu. ft. per horse- 
power hour. With a velocity of 150 ft. per second past the inlet 

120 
valve the inlet pipe cross-section is 4 X ™ X 144 + 150 = 

0.128 sq. in. per horse power. Good engines show usually from 
0.14 to 0.16 sq. in. per horse power for the inlet pipes and ports. 
Exhaust pipes and ports are usually of the same size as inlet 
pipes and ports, but are occasionally larger. In the Austro-Daim- 
ler engine inlet pipes are 0.126 sq. in. per horse power and exhaust 
pipes 0.18 sq. in. 

The effect of valve lift on engine capacity is shown in Fig. 117, 
which is plotted from Pomeroy's 1 test results. The engine 
was 90 X 120 mm. and had an adjustable inlet valve lift. The 

1 The Automobile Engineer, Feb., 1919, p. 44. 



VALVES AND VALVE GEARS 159 

valve diameter was 1.75 in. ; the valve areas (irdh) used in the test 
were 1.8, 1.4 and 0.7 sq. in. As indicated in the figure the b.h.p. 
was the same for the two larger areas up to 1,500 r.p.m. above 
which speed the larger area showed its superiority. The lowest 
lift showed marked inferiority. 

Tests of another engine of the same size with constant lift 
and with the valve areas kept the same as in the previous engine, 
by fitting different diameters of valves (1%6> 1%> and 1 in.), 
gave the results shown in Fig. 118. It is noteworthy that the per- 
formances of these three valves were practically identical up to 
1,600 r.p.m., above which speed the smallest valve showed inferior 
results and the intermediate size valve showed best results. It 
is seen by comparing Figs. 117 and 118 that valve area alone is not 
important; it is necessary to know the lift-diameter ratio also. 
The highest curve in Fig. 118 is obtained with a lift of 23.6 per 
cent of the diameter; in Fig. 117 the highest curve has a lift- 
diameter ratio of 18.7 per cent, but it is probable that still better 
results would have been obtained with a higher lift. 

Valve Materials. — Inlet valves in airplane engines under 
normal operation may reach temperatures of over 1,100°F.; 
exhaust valves may go to 1,600°F. or higher. The heat received 
by the head of an exhaust valve is dissipated in three ways: 
(1) by conduction down the stem to the guide, (2) by direct 
radiation from the back surface of the head and (3) by direct 
conduction from the face to the valve seat. The last of these is 
by far the most important. To be effective it is essential that 
the valve should have good metallic contact with its seat through- 
out the whole of the explosion stroke. If, through valve warping 
or the lodging of scale on the seat, there should be any leakage of 
gas past the valve, there will be rapid heating at the place where 
the leakage occurs and the valve will burn away at that place. 
Another prolific cause of valve burning is persistent preignitions 
in the cylinders; it is found that valves which stand up satis- 
factorily under normal operation fail very rapidly when per- 
sistent preignitions occur; with such preignitions the temperature 
of the exhaust valve may rise to 2,100°F. If the exhaust ports 
are so designed that the exhaust gases play directly on the neck of 
the valve this may become highly heated and may actually 
supply heat to the valve head instead of taking it away; in such a 
case overheating of the valve is likely to occur. It is also impor- 
tant that the valve guides should be efficiently water-cooled and 



160 THE AIRPLANE ENGINE 

should not project into the exhaust pocket so as to be heated 
directly by the exhaust gases. A final cause of overheating the 
exhaust valve is the use of an overrich mixture which may be 
still burning during the exhaust stroke. 

An interesting suggestion for valve cooling is the use of a hollow 
stem into which is put a small amount of mercury before plugging. 
The liquid mercury in contact with the hot center of the valve 
head is vaporized and is condensed again in the upper part 
of the stem. The mercury thus acts as a heat carrier abstracting 
from the valve head its latent heat each time it is vaporized. 
The vapor pressure of mercury at 820°F. is 50 lb. per square inch. 

The principal types of valve failure are (1) elongation of 
the stem, (2) distortion of the head, (3) cracks in the valve face, 
(4) wear of the stem, (5) wear of the foot, (6) burning of the 
head, (7) scaling and (8) breaking due to self-hardening. Elon- 
gation of the stem results either from the use of a steel of insuffi- 
cient strength at the working temperature or from overheating 
of the stem. Distortion of the head occurs usually when proper 
heat treatment has not been given to the valve forging before 
machining; in other cases unequal heating or softening under 
the action of high temperature may be the causes. Cracks come 
usually from cracks in the steel from which the valves are made; 
they are fairly common and are dangerous as they may result in 
the breaking away of a section of the valve. Wear on the valve 
stem occurs usually in rotary engines which produce a side 
pressure due to the inertia of the valve. Wear of the valve foot 
results from the hammering of the tappet or the wipe of the cam; 
it is diminished by hardening the foot or by the use of a cap. 
Burning is due to overheating. 

A steel to be satisfactory for exhaust valves in airplane engines 
should have the following properties as stated by Aitchison. 1 

1. The greatest possible strength at high temperatures. 

2. The highest possible notched bar value (resistance to impact). 

3. The capacity of being forged easily. 

4. The capacity of being manufactured free from cracks. 

5. The capacity of being easily heat-treated. 

6. The least possible tendency to scale. 

7. The ability to retain its original physical properties after repeated 
heatings for prolonged periods. 

8. Freedom from liability to harden on air cooling. 

1 The Automobile Engineer, Nov., 1920. 






VALVES AND VALVE GEARS 



161 



9. Freedom from distorting stresses after heat treating. 

10. Hardness to resist stem wear. 

11. Capacity of being hardened at the foot. 

12. Reasonable ease of machining. 

The best steels for exhaust valves are in five classes: 

1. Tungsten steel with not less than 14 per cent tungsten and about 0.6 per 
cent carbon. 

2. High chromium steels (stainless steel) with about 13 per cent chromium 
and about 0.35 per cent carbon. 




650 



150 850 950 

Tempera+ure , Deoj. C- 



Fig. 1 19. — Resistance of valve steels to scaling. 

3. Steel containing from 7 to 12 per cent chromium and about 0.6 per cent 
carbon. 

4. Steels containing about 3 per cent nickel. 

5. Ordinary nickel -chromium steels. 



Of these steels the first four are superior to the last. The 

nickel-chromium steels are difficult to manufacture free from 

flaws, they tend to harden during the running of the engine, they 

scale rapidly and they show no superiority at high temperatures 

over the other steels. The relative resistances to scaling are 

shown in Fig. 119, from which it is apparent that stainless steel is 

superior to the others. The tensile strengths of these steels at 

higher temperatures are given in the following table, 
li 



162 THE AIRPLANE ENGINE 

Ultimate Strength of Valve Steels, Pounds per Square Inch 



Steel 



Temperature, degrees 
Fahrenheit 




High tungsten, high carbon 39 , 600 

High tungsten, low carbon 34 , 700 

High chromium, high carbon 33 , 800 

High chromium, low carbon 27, 100 

Low chromium, high carbon 41,400 

Low chromium, low carbon 38 , 000 

3 per cent nickel, high carbon 25 , 800 

3 per cent nickel, low carbon 21,000 

Nickel chromium 23 , 500 



19,700 
14,100 
16,800 
10,700 
16,800 
15,800 
10,100 
8,700 
10,100 



The 3 per cent nickel steel is much cheaper than the others but 
is markedly inferior in tensile strength at high temperatures and 
consequently should be used only on inlet valves or for the 
exhaust valves of rotary engines. The high chromium (stainless) 
steel is highly resistant to scaling and, if of low carbon content, 
is readily machined but is not easy to forge and is liable to 
cracks. High tungsten steel retains its strength best of any steel 
at high temperatures and is fairly resistant to scaling. Exhaust 
valves which are liable to be subjected to unusually high tempera- 
tures should be of tungsten steel; for more moderate tempera- 
tures stainless steel will be more durable. Monel-metal valves 
have been used, and although they have stood up well under test 
on the Hispano-Suiza engine, they have failed rapidly on the 
Liberty engine. 

Valve Operation. — The valves of modern airplane engines are 
mechanically operated; automatic action, which is found in some 
automobile engines and in a few of the earlier airplane engines, 
must always result in lowered volumetric efficiency and capacity. 
Actuation of the valves is by means of cams acting either directly 
or indirectly. The camshafts may be placed near the base of the 
cylinders and operate the valves through push rods and rocker 
arms, or overhead camshafts may be used acting on the valves 
directly or through rocker arms. 



VALVES AND VALVE GEARS 



163 



A good example of push-rod operation is shown in the Benz 
230 engine (Fig. 78), which has separate camshafts for the inlet 
and exhaust valves, both located in the crankcase; a similar 
arrangement is used in the May bach engine (Fig. 81). In Vee 
engines the usual practice, where push rods are employed, is to 
have a single camshaft, located in the angle of the Vee, inside the 
crankcase, carrying inlet and exhaust cams for both rows of 
cylinders; the Curtiss OX and V2 engines (Figs. 60 and 62) 
show this arrangement, which is also used on the Benz 300 (Fig. 
13i). In recent years the tendency has been to do away with 
push rods and to use overhead camshafts. This last arrangement 
reduces the weight and complexity of the valve gear, and, in 
consequence of the smaller number of joints involved, makes for 
better maintenance of the valve timing. There may be either 
(1) one camshaft over each row of cylinders acting directly on 
the valves as in the Hispano-Suiza (Fig. 51), or (2) one camshaft 
acting directly on one set of valves and indirectly through a 
rocker arm on the other set as in the Siddeley "Puma" (Fig. 8), or 
(3) one camshaft acting indirectly through rocker arms on both 
sets of valves as in the Liberty (Fig. 47), Basse-Selve (Fig. 129), 
Bugatti (Fig. 67), Fiat (Fig. 76), Rolls-Royce (Fig. 70), Mercedes, 
Lorraine-Dietrich, Renault, Austro-Daimler, or (4) two camshafts 
acting directly on the two sets of valves as in the Curtiss K 
(Fig. 56) and Napier "lion" (Fig. 73). 

Cams. — The shape of the cam depends on the desired valve 
movement and on the form and location of the cam follower. 



a 



I 




Fig. 120. — Individual cam. 



Fig. 



(b) (c) 

121. — Cam forms. 



The cams are usually integral with the camshaft, which gives 
maximum security and accuracy of location, but sometimes they 
are fastened by taper pins to the hollow camshaft, as in Fig. 120; 
this arrangement permits more satisfactory hardening of the cams 
and the replacement of a worn cam but is less secure and may 
become slack. The cams are sometimes made with convex 
flanks as in Fig. 1216, or with flat surfaces tangential to circular 



164 



THE AIRPLANE ENGINE 



arcs as in Fig. 121a, or with flanks that change from concave to 
convex and a top which is concentric with the camshaft as in 
the constant acceleration cam of Fig. 121c. The cam follower 
may be flat as in Fig. 122a, rounded as in Fig. 1226, or a roller 
as in Fig. 122c, and it may be fixed on the end of the valve 
plunger or it may be mounted on a radius rod as in Fig. 122d. 
With a flat follower a convex flanked cam is used; tangential 
and constant acceleration cams are used with the other types of 
follower shown in Fig. 122. 

The work which a cam has to do is in three parts. (1) It 
must overcome the difference of gas pressures on the two sides of 
the valve. This pressure difference is important only in the case 
of the exhaust valve, which just previous to opening has a pressure 





(«) ( b ) ( c ) ( d ) 

Fig. 122. — Cam followers. 



of about 60 lb. per square inch (gage) on one side and atmospheric 
pressure on the other. With a valve 2J£ in. diameter the pres- 
sure difference is nearly 300 lb. at the instant of opening and falls 
rapidly to a negligible quantity. (2) The valve spring is operat- 
ing at all times to keep the valve closed; the compression on 
the spring is usually about 50 lb. when the valve is closed (see 
table 4). The cam must do work in compressing the spring. (3) 
The moving parts from the cam to the valve, including the valve 
and spring, must be accelerated and work must be done in giving 
them the necessary acceleration. The force required to acceler- 
ate these parts is determined by the design of cam and follower 
and by the masses that have to be accelerated. 

For maximum volumetric efficiency of the engine the valves 
should open promptly, should remain wide open as long as 
possible and then should close promptly. If a valve is to be 
opened in a given time (number of degrees of crankshaft rotation) 



VALVES AND VALVE GEARS 165 

the force required to accelerate the moving parts will' be kept a 
minimum by making the acceleration constant and thereby 
keeping the accelerating force constant. During the opening 
the moving parts must be first accelerated and then brought to 
rest; the deceleration is accomplished by the valve spring and, 
sometimes, if a push rod is used, by an additional spring acting 
on the push rod. Smooth action will be obtained when the 
deceleration is constant and has the same value as the accelera- 
tion. The cam does not necessarily do any work at all during 
the decelerating period. 

The acceleration and the force required to produce it are 
readily calculable. Suppose the valve to move from the closed 
to the fully open position in 60 deg. of crankshaft rotation and 
that the moving parts are accelerating uniformly for half that 
period, or 30 deg., and are decelerating uniformly for the next 
30 deg. Then at 1,500 r.p.m. the time, t, available for acquir- 

ing maximum velocity is .. rQ X ^q = oaa sec. If the lift is 

0.5 in. the distance, d, moved in this time is 0.25 in, and the 

acceleration, a, is given by d = ~at 2 } or a = 3,750 ft. per second 

per second. If the weight, w, of the moving parts is 1 lb. and 
if all of the parts move with the same velocity as the valve, 

w 
the force required to accelerate the parts will be — a = 110 lb. 

The force exerted by the cam will be greater than this by an 
amount equal to the valve spring compression and, at the instant 
of opening the exhaust valve, by the gas-pressure difference. 
With the numerical values given above the force exerted by the 
exhaust cam at the moment when the valve begins to open must 
be 110 + 300 + 50 = 460 lb. As there is always some tappet 
clearance to permit expansion of the valve stem without forcing 
the valve to lift, this maximum force occurs a short time after 
the cam has come into action. The force will diminish rapidly 
as the gas pressure in the cylinder falls but will tend to increase 
later with increasing spring compression. During the decelerat- 
ing period the spring pressure would have to be greater than 110 
lb. to bring the valve to rest in 30 deg. of crank rotation. During 
the closing of the valve the gas pressure difference is absent, the 
acceleration is due to spring action and the deceleration is 
brought about by pressure on the cam. 



166 



THE AIRPLANE ENGINE 



The forces exerted by the cam, which have just been considered, 
are the radial forces, R, acting along the push rod, or, in the case 
of overhead camshafts, at right angles to the outer end of the 
rocker arm. The actual pressure, N, between the cam and its 
follower acts normal to the surface of contact and will be greater 
than the- radial force throughout the accelerating period. The 
relation between these two forces is indicated by the triangle of 
forces in Fig. 1226. The side thrust, S t may be trouble- 
some. The quicker the opening of the valve the greater will be 
the acceleration force, R, and the greater will be the ratio of both 
N and S to R. In other words, the normal pressure and the 
side thrust increase much more rapidly than the radial force. 
The side thrust is particularly objectionable with valve plunger 
guides as in Figs. 1226 and c; with the arrange- 
ment of Fig. 122d, or with the cam operating 
directly on the rocker lever, a rapid opening of 
the valve can be obtained without trouble 
from side thrust. 

A good example of the constant acceleration 
type of cam with roller follower is shown in the 
Maybach engine, Fig. 123. The displacement, 
velocity, and acceleration curves both for inlet 
opening and for exhaust closure are given in 
Fig. 124; it will be seen that the valves open 
and close rapidly, and remain full open for 
considerable periods of time. The velocity 
of the exhaust valve when closing increases 
uniformly for 60 deg. of crank rotation, then 
decreases but not quite uniformly for the 
next 46 deg.; the inlet valve on opening has 
acceleration which increases for about 48 deg., 
when maximum velocity is attained, and then 
comes to rest after 60 deg. more of uniform 
deceleration. Valve displacements for the 
whole cycle are shown in Fig. 125; it will be seen that the valves 
are wide open for considerable fractions of the stroke. 

The tangential cam is used in the Liberty engine (Fig. 130) 
operating on a roller at the end of the rocker arm. With this 
type of cam, the center of curvature of the highest part of the 
cam cannot coincide with the center of the camshaft (as in the 
constant acceleration cam), and consequently the valve cannot 




Fig. 123. — May 
bach valve gear. 



VALVES AND VALVE GEARS 



167 



stay at its wide-open position. The actual valve lifts are shown 
in Fig. 126 plotted against crank position, and in Fig. 127 
plotted against piston position. The valve opening is not so 
good as in the Maybach engine, but the forces required to acceler- 



Exhaust DISPLACEMENT CURVE I nle + 

Crankshaft Degrees 
120 110 100 90 80 70 60 50 40 30 20 10 10 20 50 40 50 60 70 80 90 100 110 120 130 140 150 





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Crankshaft Degrees 

Fig. 124. — Displacement, velocity and acceleration curves for the valves of the 

Maybach engine. 

ate the moving parts are less in consequence of the longer time 
available for opening or closing the valve; with symmetrical 
cams this time is one-half the total time the valve is open. In 
the Maybach engine the intake valve is open for 223 deg. of 
crank rotation and the valve, while opening, is accelerating for 



L'xt crirsi' 



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240 220 180 140 120 110 lOO 90 80 70 60 50 4030 3040 50 60 70 80 90 100 110 120 130 140 180 220 240 250 
Diagram of Exhaust and Inlet Port Opening in Relation -to Piston Position 

Fig. 125. — Valve openings of Maybach engine. 



29 deg.; in the Liberty engine the inlet valve is open for 215 deg. 
and the valve, while opening, is accelerating for 54 deg. of crank 
rotation. As the acceleration is inversely as the square of the 
time taken to lift the valve through a given distance, the force 



168 



THE AIRPLANE ENGINE 



required to overcome the inertia of the moving valve parts would 
be 3.5 times as great for an engine using 29 deg. of crank rotation 
for the valve acceleration as for the same engine, with the same 
revolutions per minute, using 54 deg. of crank rotation. 

Valve Springs. — The function of the valve spring is to deceler- 
ate the valve moving parts during the latter half of the valve 
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Fig. 126. — Valve lifts of Liberty engine plotted against crank positions. 



closure. In addition, the exhaust valve spring must be strong 
enough to keep the valve closed during the suction stroke when 
the engine is idling. During idling the pressure in the cylinder 
may be 10 lb. per square inch below atmospheric pressure, which, 
with a 23^>-in. valve, gives a pressure difference of 50 lb. between 
the front and back of the valve. The spring pressure required 
for acceleration is normally in excess of that required to keep the 
exhaust closed during idling. 



VALVES AND VALVE GEARS 



169 



Cylindrical helical springs are employed almost universally, 
but occasionally conical helical springs may be used (Fig. 133), or, 
when minimum height is required, the rat-trap type of spring, as 
in the Curtiss engine (Fig. 128). In the larger engines two con- 
centric cylindrical helical springs are used as in the Liberty engine 



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120 200 240 270 300 330 30 60 90 120 150 20 240 
Diagram of Exhaust and Inlet Port Opening in Return to Piston Position 

Fig. 127. — Valve lifts of Liberty engine plotted against piston positions. 

(Fig. 130) and the Benz 300 inlet (Fig. 131), or even three 
springs as in the Bugatti exhaust valve (Fig. 67). An advan- 
tage of multiple springs is that, in case of breakage of one spring, 
the valve cannot fall into the cylinder. 




Fig. 128. — "Rat trap" valve spring (Curtiss engine). 

The maximum safe working load, P, in pounds, on a cylindrical 
helical spring of outside diameter D in. and with steel of diameter 
d in. is given by 



170 



THE AIRPLANE ENGINE 



where S is the safe shearing stress of the material in pounds 
per square inch. The value of $ varies from about 80,000 to 
150,000 (increasing as the diameter of the wire diminishes) for 
springs that are used intermittently. For continuous use, as in 
an airplane engine, about half these values should be employed, 
or, for the usual sizes of spring steel, about 60,000 lb. 

The deflection, /, in inches of one coil of a cylindrical helical 
spring is given by 

8nPD* 
j d*G 

where G is the modulus of elasticity in shear and may be taken as 
12,000,000 lb. per square inch. The following table gives safe 
loads, P, and the corresponding deflection,/, of one coil for various 
springs. It is calculated for S = 60,000 lb. per square inch. 

Safe Loads and Deflection of Cylindrical Helical Springs 



Wire gage 


d in. 


Outside diameter of spring, D in. 




1.75 


2.00 


2.25 


2.50 


No. 8 


0.162 
0.192 
0.205 
0.225 

0.242 


P 

f 
P 

f 
P 

f 
P 

f 
P 

f 


64. 

0.23 
107. 

0.20 
131. 

0.18 
175. 

0.16 
225. 

0.11 


55.5 

0.33 
92.5 

0.26 
113. 

0.25 
150. 

0.22 
195. 

0.16 


48.8 

0.42 
81. 

0.34 
99. 

0.32 
132. 

0.29 
170. 

0.20 


43.5 


No. 6 


0.57 

72. 


No. 5 


0.42 

88.5 


No. 4 


0.36 
118. 


No. 3 


0.36 
152. 




0.30 



For square steel of side d in. the tabular valves of P should be 
multiplied by 1.2, and the values of/ by 0.59. 

The valve spring retainer is usually a washer or cupped disc 
with a downwardly turned flange to center the spring. The 
retainer is fastened to the valve stem in various ways. It is 
sometimes held by a nut which screws on to the threaded upper 
end of the valve stem and is locked by a split pin as in the Basse- 
Selve engine (Fig. 129); or it is held by a cotter through the 
valve stem, locked in position by wire clips as in the Maybach 
engine (Fig. 123); or the valve stem is turned to a smaller diame- 
ter for a short length near the top and held by a conical split 



VALVES AND VALVE GEARS 



171 



collar as in the Liberty engine (Fig. 130), the Benz 300 (Fig. 
131) and the Fiat engine (Fig. 134). In those engines in which 
the cam acts directly on a flat cam follower on top of the valve 
stem this follower may serve to hold the spring retainer. In 
the Hispano-Suiza engine (Fig. 132) the upper end of the valve 
stem is slotted to receive the disc-shaped retainer and is threaded 
internally. The retainer slips over the valve stem and is pro- 
vided with a key which fits into the slot and prevents rotation; 
it has fine notches on its upper surface to mesh with correspond- 
ing notches on the lower face of the follower. The follower stem 
screws into the hollow valve stem and after being screwed into 
the desired position the retainer is permitted to come into contact 




Fig. 129. — Valve gear of Basse-Selve engine. 



with it and locks it in position. A similar arrangement is shown 
for the Siddeley "Puma" engine (Fig. 133), but in this case a 
volute steel spring of rectangular section is used. 

With dual valves one spring, or a pair of concentric springs, 
can be made to serve for the two valves as in the Fiat engine 
(Fig. 134). The springs are in a steel yoke or cage. The inner 
spring is mounted on a central guide tube while the outer spring 
is retained by the yoke. The lugs on the two sides of the yoke 
fit over split locking cones which are held by a spring ring against 
the grooves turned in the upper ends of the valve stems. 

Rocker arms are usually pivoted in plain bearings except 
in German practice, where ball or roller bearings are commonly 



172 



THE AIRPLANE ENGINE 



used, as in the Basse-Selve engine (Fig. 129) and the Benz 300 
(Fig. 131). In this last 
engine, which has one in- 
take and two exhaust 
valves, the rocker for the 
intake valve is brought 
obliquely under the 
double rocker of the ex- 
haust valves and the push 
rods are thereby kept in 
line and close to the cyl- 
inders without unduly 
shortening the length of 
the rocker. Double 
springs of very low height 
are required for the in- 
take valve and are sunk 
in the dished cylinder 
head. 





Fig. 130. — Valve of Liberty 
engine. 



Fig. 131.— Valve gear of Benz 300. 



Adjustment has to be provided for the tappet clearance. A 
common method is that shown for the Hall Scott A7A engine 



VALVES AND VALVE GEARS 



173 



(Fig. 135), in which the hardened steel set screw at the end of the 
rocker arm is clamped in position by a lock screw at the end of 
the split rocker arm; in this engine the clearance is 0.02 in. when 
the engine is cold. In the Liberty engine (Fig. 130) the adjusting 




Fig. 132. — Valves of Hispano-Suiza engine. 

screw is held in place by a lock nut. The amount of clearance 
should exceed the expansion of the valve stem and depends 
mainly on the length of the valve stem. As the inlet valve is 
colder than the exhaust valve it requires less tappet clearance; 
in the Liberty engine the inlet clearance is 0.015 in., the exhaust 




Fig. 133. — Valves of Siddeley "Puma" engine. 

0.020 in. Too much clearance is to be avoided, as producing 
noise and possible breakage of parts. • 

Valve Timing. — If the inlet and exhaust valve openings could 
be made unobstructed and large enough, and if the openings 
and closings were sufficiently rapid, each valve could be timed to 



174 



THE AIRPLANE ENGINE 



open at the beginning of a piston stroke and to close at the end 
of that stroke. In actual engines it is necessary for the valves to 
depart from this timing in the interests of high volumetric effi- 
ciency and capacity. Especially must the exhaust valve open 
early and the inlet valve close late. The time of opening of the 




Fig. 134. — Valves of Fiat engine. 

inlet valve is related to the time of closing of the exhaust valve: 
the exhaust valve usually closes completely shortly after the 
inlet valve starts to lift. The valve does not start to open until 
the tappet has moved a distance equal to the tappet clearance. 
The inlet valve opening is usually 10 to 15 deg. past top dead 
center (T.D.C.) but sometimes occurs before top dead center. 




Fig. 135. — Tappet adjustment, Hall-Scott engine. 

Inlet valve closure is usually about 40 deg. past bottom dead 
center (B.D.C.), which corresponds to about 10 per cent of the 
return stroke of the pistpn and causes a corresponding decrease 
in the compression ratio. The exhaust valve opens from 45 to 
50 deg. before bottom dead center and closes about 10 deg. past 
top dead center. The actual timing for any engine should be 



VALVES AND VALVE GEARS 



175 



determined by operation of the engine and observation of the 
timing giving maximum capacity. The inlet valve should not be 
opened until a partial vacuum is established in the cylinder; too 
late a closure will result in reduction of compression below that 
obtainable with proper timing. Valve timings for various 
engines are given in the following table: 



Examples of Valve Timing 



Engine 



Inlet 


Opens 


Closes 


15°L 


40°L 


15°L 


45°L 


12°L 


40°L 


10°L 


45°L 


TDC 


37°L 


TDC 


45°L 


10°L 


52H°L 


10°E 


62°L 


TDC 


40°L 


2°L 


51°L 


8°E 


35°L 


5°L 


45°L 


10°-15°L 


35°-50°L 



Exhaust 



Opens 



Closes 



Duration of opening 



Intake Exhaust 



Hall Scott — A7a. . . 
Hall Scott— A5a. . . 

Curtiss— 90 

Liberty 

Curtiss— K12 

King-Bugatti 

Hispano-Suiza — 180 
Hispano-Suiza — 300 

Mercedes — 180 

Mercedes — 240 

Maybach— 300 

Benz — 200 

Average 



45°E 

50°E 

45°E 

50°E 

47°E 

47H°E 

48°E 

62M°E 

40°E 

52°E 

33°E 

55°E 

45°E 



10°L 
10°L 

TDC 
10°L 

TDC 

uy 2 °L 

10°L 

29H°L 

10°L 

16H L 
7°L 
18°L 
10°L 



205° 
210° 
208° 
215° 
217° 
225° 
2223-2° 
243° 
220° 
229° 
223° 
230° 
200°-220° 



235° 

240° 

225° 

240° 

227° 

245° 

238° 

272° 

230° 

248K' 

220° 

253° 

235° 



early. L = late. TDC = top dead center. 



CHAPTER VIII 
RADIAL AND ROTARY ENGINES 

Radial and rotary engines are characterized by having the 
cylinders disposed at equal angular intervals around a complete 
circle. The number of cylinders may range from 3 to 20 or 
more. It is not possible to arrange more than 10 or 11 cylinders 
in a circle without increasing the size of the crankcase to dimen- 
sions which give an over-all diameter too large for use in an 
airplane. If more cylinders are desired they have to be arranged 
in two planes or banks, with an equal number of cylinders in 
each plane; with air cooling, the cylinders of the rear row are 
staggered with reference to those of the front row; with water 
cooling, they may lie exactly behind those of the front row. 

Fixed-radial engines have stationary cylinders and a revolving 
crankshaft; there are usually as many cranks as there are rows of 
cylinders, although two cranks have been used with a single row 
of cylinders. 

Rotary engines have rotating cylinders and a fixed crank- 
shaft; in this case the propellor hub is attached to the rotating 
crankcase. 

Double -rotary engines have both cylinders and crankshaft 
rotating but in opposite directions. With this type, two arrange- 
ments are possible for utilizing the power developed: (a) two 
propellers may be used mounted on the crankcase and the crank- 
shaft respectively and therefore rotating in opposite directions 
but with right- and left-hand pitches respectively, so that both 
give thrust in the same direction; (b) the crankcase may be geared 
to the crankshaft (with reversal of motion) and the power 
absorbed by a single propeller on the crankshaft or crankcase. 

Air Cooling. — The disposition of the cylinders of a radial or 
rotary engine, in a plane at right angles to the wind and the slip 
stream, gives these types a unique opportunity for direct air 
cooling of the cylinders. This is especially the case with rotary 
engines, which churn through the air as well as meeting the 
incoming wind. 

With cylinders in line, as in multicylinder vertical or Vee 
engines, air cooling is practicable only with a blower for supplying 

176 



RADIAL AND ROTARY ENGINES 177 

the cooling air and with a system of ducts for distributing the 
air to the different cylinders. Renault Freres have built engines 
of this type up to 12-cylinder Vees but there are considerable 
difficulties in obtaining high mean effective pressures and low fuel 
consumption with such cooling as can be obtained in this manner. 
This type has not met with general favor. 

Radial engines are often water-cooled although this practice 
sacrifices one of the most important potential advantages of 
the engine. Rotary engines are always air-cooled and could not 
be readily water-cooled, both as a result of the mechanical diffi- 
culties in getting water to and from the rotating cylinders, and 
also because of the excessive centrifugal stresses which would be 
set up in the connections between the cylinder and the crankcase 
as a result of the increased mass of the cylinder with its water and 
jacket. 

Advantages of Radial and Rotary Engines. — The primary 
advantage of radial and rotary engines over other types is in 
small weight per horse power. This results from two principal 
causes. 

1. The engine can be air-cooled, and the water pump, water, 
jackets, water pipes and radiator eliminated. The only impor- 
tant additional weight is that of the cooling fins on the cylinder. 

2. The crankcase and crankshaft are much shortened and 
lightened; the big ends of the connecting rods are also lighter. 

The net result of these reductions is a decrease of almost 40 per 
cent in weight per horse power of the power plant. The Liberty 
engine with water and radiator weighs about 2.6 lb. per horse 
power; large air-cooled fixed radial engines have a weight of 
about 1.6 lb. per horse power. 

Other advantages of the radial and rotary engines are short 
over-all length, which permits better location of gasoline tanks, 
pilots, etc.; immediate accessibility of the cylinder heads, and 
easy accessibility of the rest of the engine on removal of the 
cowling; ease of mounting on and detaching from the fuselage, 
the attachment being to a vertical plate ; lowered air resistance in 
small sizes which can be accommodated without special enlarge- 
ment of the fuselage — in larger powers it is necessary to increase 
the size of the front of the fuselage and thereby increase its resis- 
tance until it is likely to be as great as that of a water-cooled 
engine and its radiator; engine balance as good as in the best 
arrangements of vertical and Vee engines. 

12 



178 THE AIRPLANE ENGINE 

Disadvantages. — The principal disadvantage of the air- 
cooled radial and rotary engines up to the present time has been 
a lower m.e.p. and higher fuel consumption than in water- 
cooled engines. These features are especially true of the rotary 
engine, which suffers the further disability that its revolutions per 
minute must be low in order to keep down the air-churning 
resistance and the centrifugal forces exerted by the cylinders 
on the crankcase. With improved constructions of cylinders 
(see p. 202) it is probable that the performance of fixed air- 
cooled radial engines will not be markedly different from that 
of water-cooled engines. Other disadvantages are larger oil 
consumption, which is particularly marked in rotary types but 
can be kept down in fixed radials ; large over-all diameter neces- 
sary for large powers, requiring an enlarged fuselage and possibly 
limiting the power developable per unit; large crankpin pressures 
in fixed radials; large gyroscopic effect in rotaries, affecting the 
maneuvering qualities of the plane; large inertia effect in rotaries, 
retarding the speeding up of the engine on opening the throttle 
but giving more uniform speed at low speeds or with missing 
cylinders. 

Number of Cylinders and Firing Order. — With a single crank, 
the number of cylinders of a radial or rotary engine must be 
odd; the firing order will follow the direction of rotation of the 
crank in fixed radials and will be in the opposite direction to 
rotation of the cylinders in rotaries. In both cases it will skip 
alternate cylinders and will have occurred in all the cylinders in 
two complete revolutions of the crank or cylinders. If the cylin- 
ders are numbered serially the firing order will be 1, 3, 5, 7 . . . 
2, 4, 6, 8 . . . 

With a two-throw crank and equal angular spacing of the 
cylinders the number of cylinders acting on each crank should be 
odd; the total number of cylinders will be even. (The Smith 10- 
cylinder engine 1 with one row of cylinders has a two-throw crank 
so that it is really equivalent to two rows of five cylinders each.) 
The firing should occur alternately on the two cranks which are 
at 180 deg. The firing order for regular impulses for a 10- 
cylinder engine, with serial numbering, will be 1, 8, 5, 2, 9, 6, 3, 10, 
7, 4. The 20-cylinder Anzani air-cooled rotary has four rows of 
five cylinders each. This arrangement keeps the over-all diameter 
small but impairs the cooling of the two rear rows. 

1 Jour. S. A. E., Jan., 1919. 



RADIAL AND ROTARY ENGINES 



179 



With a two-throw crank and equal spacing of cylinders in 
each row but with the cylinders of the second row immediately 
behind those of the first row 1 the number of cylinders in each row 
may be odd or even. Firing is alternately from front to back row 
cylinders except when there is an even number of cylinders in 
each row. In that case two successive impulses will come on one 
crank at the end of each revolution but the angle between 
impulses will be constant. 

Rotary Engines. — In a rotary engine (Fig. 136) the cylinders 
rotate about the crankshaft as a center; the pistons rotate about 




Fig. 136. — Diagram of rotary engine. 

the crankpin as a center. The angular velocity of the pistons 
about the crankshaft as center is constant since it must be the 
same as that of the cylinders which are rotating at uniform veloc- 
ity; the angular velocity of the pistons about the crankpin as 
center would be constant only when the connecting rods were 
infinitely long or the crank throw infinitesimally small. The 
result of the rotations about the two centers is to give the piston 
a reciprocating motion relative to the cylinders as shown in 
Fig. 136 (an analysis of this motion is given on p. 507). As the 
cylinders rotate the pistons will assume the positions shown rela- 
tive to the cylinders. If the crank is fixed with its throw verti- 
cally upward the piston will always be at the in dead center 
1 See 20-cylinder water-cooled Anzani, Aviation, Feb. 15, 1920.J 



180 THE AIRPLANE ENGINE 

when the cylinder reaches the position 1; it will be at the out 
dead center for the position vertically below 1. One revolution 
of the cylinders completes one in and out stroke of each piston. 

Certain special problems arise in the construction of the rotary 
engine. One of the most important of these is the accommoda- 
tion of seven or nine connecting rods on one crankpin; this 
problem occurs also with fixed-radial engines. Another is the 
attachment of various members to the rotating cylinders. The 
exhaust manifold is always eliminated and the exhaust gases 
discharge directly past the valve to the air. The carburetor 
(or equivalent device) must be stationary and discharges into the 
crankcase; separate induction pipes to the individual cylinders 
may or may not be provided. There is no possibility of continu- 
ous circulation of lubricating oil; the oil is used up as it is fed. 
The cylinders must be machined all over to exact dimensions to 
avoid unbalanced centrifugal forces. 

Gnome Engine. — A longitudinal section of the 100-h.p. Gnome 
Monosoupape (single-valve) engine is shown in Fig. 137. The 
engine has nine cylinders (each machined out of a 6-in. solid 
nickel-steel bar) arranged at equal angular distances around the 
crankcase. The bore is 110 mm., the stroke 150 mm., clearance 
volume 365 cu. cm., normal speed 1,200 r.p.m., weight 260 lb. 

The cycle of operations is as follows : starting with any cylinder 
on top center and the exhaust valve open, the cylinder draws in 
air through the exhaust valve until its closure at 45 deg. before the 
bottom center; that is, air is drawn in for 135 deg. rotation of 
the crank. During the next 25 deg. a vacuum is created in the 
cylinder until, at 20 deg. from the bottom center, the admission 
ports are uncovered and a rich mixture enters from the crankcase, 
mixing with the air already in the cylinder and forming an explo- 
sive charge. The ports are again covered at 20 deg. past bottom 
center and, as the cylinder rotates to its top center, compression 
occurs. Ignition takes place at 20 deg. before top center and 
the cylinder rotates on its power stroke until it is 90 deg. past 
top center, when the exhaust valve opens and remains open for 
the following 405 deg. rotation. It will be seen that the exploded 
charge is expanded for only half the stroke and consequently there 
is no possibility of high capacity or efficiency with this engine. 
One important reason for the early exhaust of the exploded 
mixture is to prevent overheating of the cylinders. It is found 
that a late exhaust will cause overheating. 



RADIAL AND ROTARY ENGINES 



181 




182 THE AIRPLANE ENGINE 

The mixture in the crankcase is at all times too rich to be 
explosive. Air is drawn in through the open rear end of the 
hollow stationary crankshaft. The fuel is supplied under air 
pressure and sprays from the fuel nozzle into the crankcase, 
where it is churned up with the air. There is no throttle valve, 
but the fuel supply can be controlled by a regulating valve to 
adjust the mixture strength. The power output of the engine can 
be varied through a small range only, by the use of the regulating 
valve or switching the ignition on and off. Variation of the 
power by cutting out the ignition on one or more cylinders gives 
unequal impulses, and causes fouling of the cylinders which are 
not firing. 

The magneto is mounted on the face of the back plate remote 
from the engine. It is driven through a spur gear which meshes 
with the large gear keyed to the thrust box casing. The gear 
ratio is 4 :9, that is, the magneto armature makes nine revolutions 
to four of the engine, and as the magneto gives two sparks per 
revolution there will be nine sparks in two revolutions of the 
engine. The current from the magneto goes to the distributor 
brush, which makes contact in turn with the nine metal segments 
of the distributor ring. The distributor ring revolves with the 
engine and consequently 18 contacts are made in two revolutions 
of the engine, but no spark passes on the exhaust stroke as none 
is generated at the magneto. 

The air pump (for the fuel pressure) and oil pump are mounted 
on the back plate and driven from the same large gear as the 
magneto. The oil pump delivers oil into two pipes of equal 
size. Of the oil going into one pipe about one-third flows through 
a branch into the thrust box, oiling the thrust ball race and main 
engine ball race. The surplus oil overflows into the crankcase 
through holes drilled for this purpose, and passes on to the 
cylinder walls through the ports in the base of the cylinder. The 
main supply of oil passes up the big crank web through a hollow 
plug in the center of the hollow crankpin, down the short end 
crank web into the hollow short end of the crankshaft, 
whence it is conveyed by a series of holes to the cam pack. The 
oil then passes through grooves between the cams and is thrown 
centrifugally over the interior of the cam box, lubricating the 
cams, cam rollers, tappets, planet gear wheels, and the cam 
box and nose-piece ball races. The oil then overflows back into 
the crankcase and passes on to the cylinder walls as in the case 



RADIAL AND ROTARY ENGINES 183 

of the overflow from the thrust box. Some of the oil also passes 
along the hollow tappet rods to the rocker arm pins. 

The oil from the other pipe flows up inside the long end crank 
web into an annular space around the brass plug in the long end 
crankpin and out of holes in its balls to two grooves or channels 
cut in the ends of the bore of the master connecting rod big end. 
Holes are drilled from these grooves to each wristpin and the 
wristpins are drilled to correspond with these holes, so that the 
oil may pass through to lubricate the wristpin bushings. From 
these the oil passes into steel tubes (which are fixed to the con- 
necting rods) and along the tubes, oiling the gudgeon or wristpins 
and bushes. In later type engines the steel tubes are dispensed 
with, and the oil passes along the face of the connecting rods to 
the gudgeon pins and bushes. The overflow from the gudgeon 
pins passes through holes in the side of the pistons, and lubricates 
the rings and the cylinder walls. The surplus oil is blown out 
through the exhaust valve and lubricates the exhaust valve-guide 
and stem. 

The crankcase is made of two steel stampings bolted together 
by steel bolts and centered by dowel pins. The nine cylinders 
are each gripped tightly by the two parts of the crankcase and 
prevented from turning by a small key. The crankcase is 
not directly supported on the crankshaft, but carries on its faces 
plates or covers, known respectively as the cam box and the 
thrust box. The thrust box contains the main ball race and a 
self-aligning double-thrust race. The cam box contains the 
planet gears and the cam pack which actuates the exhaust 
valves, and one radial ball race. The nose piece which carries 
the propeller is mounted on the cam box. 

The pistons are of cast iron with concave heads. A portion 
of the trailing edge is cut away to allow the piston in the adjoin- 
ing cylinder to clear. Each piston is fitted with an obturator 
ring about 0.6 mm. thick in a groove around its top. This 
obturator ring is of cupped form and is pressed out against the 
cylinder wall by the gas pressure, thus preventing leakage past 
the piston. A packing ring is fitted behind the obturator ring 
and in the same groove. A wipe ring which is made of cast 
iron is also fitted in a groove situated just below the obturator 
ring. The piston is fastened to its connecting rod by means of a 
hollow steel gudgeon pin fixed in lugs on the underside of the 
piston head by means of a tapered set screw. 



184 THE AIRPLANE ENGINE 

Connecting Rods. — The connecting rod assembly consists 
of a master connecting rod, to which eight auxiliary connecting 
rods are attached by means of wristpins. All the rods are of 
H-section and the auxiliary rods are bushed at both ends with 
phosphor-bronze bushes. The master connecting rod big end runs 
on two ball bearings; its small end is bushed with phosphor bronze. 

The single valve in the cylinder head performs the following 
functions: (a) It acts as an exhaust valve; while so doing its 
temperature is raised; (6) it admits to the cylinder a quantity 
of air sufficient for combustion of the charge entering later 
through the ports at the base of the cylinder. During this 
portion of the cycle it is cooled. 

The valve is 60 mm. diameter and has a lift of 10.5 mm. It 
is mounted in a steel cage which also carries the rocker arm 
fulcrum pin, and is mechanically operated by means of a hollow 
steel tappet rod and steel rocker arm. The valve stem slides 
in a cast-iron bush at the center of the cage which is held in 
position by means of a locking ring screwed into the cylinder 
head. The valve is made heavier than is necessary for mechani- 
cal strength and is of such weight as to balance the centrifugal 
action of the tappet rod which would otherwise tend to keep the 
valve open. The valve spring is spiral and encircles the valve 
stem, taking its bearing against the valve cage and a detachable 
collar on the valve stem. The valves are operated by the cam 
pack, which consists of nine cams keyed to a bronze-bushed sleeve 
rotating on the small end of the crankshaft. The cams operate 
the tappet rods which work the overhead rocker arms. Each 
tappet rod is formed of a tappet and a rod jointed together. 
The tappet works in a guide in the cam box, and at its inner end 
is a roller which bears against the cam. The tappet rod extends 
from the joint to the rocker arm of the exhaust valve, and is 
adjustable. The cam pack is driven at half the engine speed by 
planet gears, which are fitted on the inner face of the cover of the 
cam box. The engine is running at twice the speed of the cam 
pack, so that the rollers at the bases of the tappet rods are over- 
taking the cam pack. The clearance between the rocker arm 
and the bottom of the slot in the valve stem, when the tappet 
roller is at the bottom of the cam, should be 0.5 mm. when the 
engine is cold. In later type engines the rocker arm engages 
the valve stem by means of a roller which bears against the end 
of the stem. 



RADIAL AND ROTARY ENGINES 



185 



Le Rhone. — The nine-cylinder 110-h.p. Le Rhone engine 
is shown in Figs. 138 and 139. The bore is 112 mm., stroke 170 
mm., the normal speed is 1,200 r.p.m., weight 323 lb. The 80- 
h.p. nine-cylinder engine is of 105 mm. bore, 140 mm. stroke. 
The cylinder has a cast-iron liner. This engine differs from the 
Gnome engine in several important respects. It has an inlet 
valve as well as an exhaust valve and consequently has to be 



■ Valve rocker gear 




D= Inlet cam 

E= Exhaust cam 

F= Internal gear 

mountea eccentrically 



Fig. 138. — Transverse section of Le Rhone 110. 



provided with inlet pipes from the crankcase to each inlet valve 
cage. The valve timing is as follows: the exhaust valve closes 
5 deg. after top center on the suction stroke; the inlet valve opens 
13 deg. later or 18 deg. after top center; the inlet valve closes 
35 deg. after bottom center and compression goes on till 26 deg. 
before top center, when ignition occurs. The expansion occupies 
125 deg. of the power stroke when the exhaust valve opens at 
55 deg. before bottom center and remains open for 140 deg. 



186 



THE AIRPLANE ENGINE 



This timing differs notably from that of the Gnome engine and 
gives much more complete expansion of the gases. 

A rudimentary carburetor (see p. 288) is located at the rear 
end of the hollow crankshaft, admitting an explosive mixture 
to the crankcase. Control of power output is through the 
throttle valve. Ignition is by a magneto and distributor similar 
in location and general construction to those of the Gnome 



Rocking lever fulcrum 






SpTjSd^ i 


\ 









Fig. 139. — Longitudinal section of Le Rhone 110. 



engine. The pistons are convex and of semi-steel. The con- 
necting rods are of the slipper type (see p. 204). 

The two valves on each cylinder are actuated by the motion 
of a rocking lever which is fulcrumed at its middle in ball bearings. 
This lever is operated by a valve-actuating rod which receives its 
motion from the trailing end of a cam-follower lever. The cam- 
follower lever is fulcrumed at its middle and carries an inlet 
valve cam follower at the forward end and an exhaust valve cam 



RADIAL AND ROTARY ENGINES 



187 



follower at the trailing end. The two cam-followers are in 
different planes and are acutated by the inlet and exhaust cams 
respectively. These cams are lobed plates mounted on a spider 
running in ball bearings on a shaft eccentric to the crankshaft. 
The spider carries an internal gear into which meshes an external 
gear mounted in ball bearings on the crankshaft and rotating 
with the crankcase. These arrangements are shown in the right 
half of Fig. 138 and also in Fig. 139. The external gear has 45 
teeth; the internal gear 50 teeth; consequently one complete 
revolution of the engine will produce nine-tenths of a revolution 
of the cams. The engine is overrunning the cams and would 

100 
90 
80 
70 



v 60 

5 















IJ-~ 
















6^^^- 


















" | 






















Gross M. 


E.P. 






\l 
























































































































windage __ 
j=-| 




| 









95 
90 
85 ex.; 
80 5= 
75 
70 



800 



900 1000 1100 1200 

Revolutions per Minute 



1300 



Fig. 140. — Performance curves of Le Rhone 80. 

overrun them one complete revolution in 10 revolutions of the 
engine. Each cam has five lobes. Each inlet and exhaust valve 
should be opened once only in two revolutions of the engine, and 
this will be accomplished when the engine overruns the cam one- 
fifth of a revolution. As the engine overruns the cams one- 
tenth of a revolution each engine revolution it is evident that 
two revolutions of the engine are required to complete the 
opening and closing of all the valves. 

The action of centrifugal force on the valve-actuating rod 
causes it to press continuously against the valve rocker lever. 
At low speeds of revolution this force may not be sufficient to 
open the exhaust valve at the desired time and consequently an 
exhaust cam is necessary to push the valve rod out. At high 
speeds the operation of both valves can be taken care of by the 



188 



THE AIRPLANE ENGINE 




O 



RADIAL AND ROTARY ENGINES 



189 



inlet cam if it is properly shaped for that purpose. The valves 
are brought back to their seats by spiral springs at low speeds; at 
high speeds centrifugal force closes the valves. 

Performance curves for the 80-h.p. Le Rhone are given in 
Fig. 140. 

Clerget. — The Clerget rotary engines are built with 7, 9 or 11 
cylinders. The 130-h.p., nine-cylinder engine shown in longi- 
tudinal section in Fig. 141 is 120 mm. bore, 160 mm. stroke, makes 



Cam Gear Box 



Exhaust Tappet 
6uide: 



%•" 

Key- way-- 

Inlet Gear 

R> n 3- ~" 

In let Cam. 
Exhaust 
dear 
Rings' 



Cam 6ear\ 
Cover..--'' 




-Inlet Tappet. 



Locking Nut. 
Exhaust Cam. 
Locking Nut 
■ON Hole. 

Iniet Cam. 



Fig. 142.— Cam gears of B.R. 2. 



1,250 r.p.m., weighs 381 lb., develops 135 h.p. and has a compres- 
sion ratio of 4. Its points of difference from the previous engines 
include the use of an aluminum piston, tubular connecting rods, 
inlet and exhaust valves operated by means of separate cams, 
tappets and rocker arms, and a double-thrust ball race which is a 
pure thrust bearing and distinct from the combined thrust and 
radial bearings of the other engines. 

The inlet and exhaust cam plates are driven at nine-eighths 
the engine speed by separate internally-toothed gears mounted 



190 



THE AIRPLANE ENGINE 



inside and keyed to the cam-gear case. These mesh with external 
gears mounted eccentrically on the crankshaft; the cams are 
attached to these external gears. This arrangement is the 
reverse of that used on the Le Rhone engine. The cam plates 
overtake the engine once in eight revolutions. Each cam plate 
has four lobes so that in eight revolutions each tappet will be 
lifted four times, or once in two revolutions. A sectioned per- 
spective view of the similar cam-gear box of the B.R.2 rotary 
engine is shown in Fig. 142. The four cams on each gear are 
simply rearward extensions of every fourth tooth. 

The valve timing differs in some respects from that of the 
Gnome and Le Rhone. At top center of the suction stroke 






no 



160 



J50 
v 

8.140 
v 



1130 



120 

















Jo^jJ- 


Z-^^~~~ 






















Effects 


ve Hp 







































1100 1150 1200 1250 1300 

Revolu+ions per Minu'te 

Fig. 143. — Performance curves of Clerget 130. 



1350 



both exhaust and inlet valves are open. The inlet opens 5 deg. 
before top center; the exhaust closes 5 deg. past top center. The 
inlet remains open till 58 deg. past bottom center (or a total of 
153 deg.) and compression begins. Ignition is at 25 deg. before 
top center and exhaust begins 68 deg. before bottom center. 
The carburetor is located at the rear end of the hollow crank- 
shaft. Fuel is injected under air pressure through a jet which is 
controlled by a needle valve. The air supply is controlled by a 
cylindrical throttle valve. Equal movement of both throttle 
lever and needle valve lever controls air supply only. Operation 
of the throttle lever alone controls both air and fuel. The 
charge entering the crank passes to the annular inlet chamber 
at the rear of the crankcase and then by the separate inlet pipes 
to the cylinders. The connecting rod assembly is similar to that 
of the Gnome engine (see p. 203). 



RADIAL AND ROTARY ENGINES 



191 



The performance curve for this engine (Fig. 143) is typical 
of rotary engines. The effective horse power goes through a 
maximum at 1,250 r.p.m. but is very flat for a considerable range 
of speed. The rapidly increasing difference between the effective 
horse power and the indicated horse power is due to the rapid 
increase in the air-churning resistance. 

B.R.2 Engine. — The British B.R.2 engine is one of the largest 
air-cooled rotary engines. In general construction it is similar 
to the Clerget. The cylinders are of aluminum with steel liners 
and steel head. The cylinder diameter is 140 mm., stroke 180 
mm., compression ratio 5.01, brake horse power 230 at 1,300 
r.p.m., weight dry 498 lb., weight per brake horse power dry 




Fig. 144.— Thrust box of B.R.2. 

2.16 lb. The cam gear for this engine is shown in Fig. 142 and 
follows exactly the same principle as the Clerget cam gear. The 
thrust box contains two ball bearings and a thrust bearing which 
differs from the Clerget in having one row of balls only. This 
single-thrust bearing is adapted both for pusher and tractor 
use as indicated in Fig. 144; a very small clearance is left for the' 
travel of the crankcase along the crankshaft when changing from 
pusher to tractor. 

Double Rotary. — The double-rotary engine has cylinders 
revolving in one direction while the crankshaft revolves in the 
other direction. The effective speed is the sum of the two speeds 
so that the power of an engine in which both cylinders and crank- 
shaft revolve at 900 r.p.m. is the same as that of a radial or 
rotary engine of the same dimensions operating at 1,800 r.p.m. 
Such a speed is permissible in radial engines but would give 



192 



THE AIRPLANE ENGINE 



excessive air-churning resistance in a rotary. There is no reason 
why even higher effective speeds — up to 2,400 r.p.m. — may not 
be practicable with this type, if the volumetric efficiency of the 
engine can be maintained and if the cylinders can be kept cool 
enough. In any case this arrangement leads to a combination 
of high engine speed and low propeller speed with consequent 
high propeller efficiency. It has important advantages over all 
other types in (a) the possibility of the elimination of unbalanced 




Carburetor 



Fig. 145. — Longitudinal section of Siemens-Halske double rotary. 

gyroscopic effects, which is an advantage for maneuvering, and 
(b) the elimination of the unbalanced turning moment exerted 
by the engine on the plane. This unbalanced turning moment 
is a constant, though small, power drag on the plane and its 
elimination is a distinct advantage. 

The only engine of this type which has been in production is 
the Siemens-Halske 11-cylinder engine (Fig. 145), which was 
brought out in 1918 and develops 200 h.p. at 900 r.p.m. of both 
cylinders and crankshaft, or a virtual speed of 1,800 r.p.m. 



RADIAL AND ROTARY ENGINES 193 

The single propeller is mounted on a nose attached to the revolv- 
ing crankease; the torque of the crankshaft is transmitted to the 
crankcase by securing a bevel wheel to the crankshaft and a 
similar gear, facing it, to the crankcase and mounting an inter- 
mediate pinion between the two on a stud which is fastened to 
the stationary cylindrical housing at the rear of the engine. 

The carburetor is mounted on a stationary hollow extension 
of the crankshaft. The combustible charge is drawn in through 
the hollowcrank shaft to the crankcase and goes from the annular 
inlet chamber at the rear of the crankcase to the individual inlet 
pipes. The inlet and exhaust valves are operated through two 
cam plates which are loose on the crankshaft and are rotated 
through double reduction gears from an internal gear attached 
to the crankcasing. The engine is supported by steel rods both 
before and behind the cylinders. 

The weight of the engine complete is 427 lb., which at 240 
maximum h.p. gives a weight of 1.78 lb. per horse power. The 
fuel economy is as good as with stationary engines and is much 
better than with other rotaries. 

Other designs of double rotaries with two propellers (right- 
and left-hand respectively) forward of the cylinders, attached 
to the crankshaft and crankcase respectively, have not passed the 
experimental stage. The efficiency of a pair of propellers close 
together but operating in opposite directions has been found to 
be but little inferior to that of a single propeller. There is 
consequently the possibility of the development of a satisfactory 
double-rotary engine on the lines indicated. 

Radial Engines. — In a fixed radial engine the cylinders are 
stationary and the crankshaft revolves. Three to eleven cylin- 
ders can be accommodated in a single row or bank, but two rows 
with a two-throw crank must be adopted if a larger number of 
cylinders is desired or if it is necessary to cut down the over-all 
diameter. The two-throw crank eliminates the need for counter- 
balance weights but increases the length of the engine and intro- 
duces difficulties in the air cooling of the rear row cylinders. 

Radial engines offer certain special construction problems. 
The most important are the balancing of the masses at the crank- 
pin and the avoidance of excessive pressures on the crankpin. 
It is possible to operate radial engines at speeds as high as those 
used in vertical and Vee engines, but special care must be taken 
to prevent the overheating of air-cooled cylinders and the over- 

13 



194 



THE AIRPLANE ENGINE 



loading of the crankpin. There is no fundamental reason why 
the mean effective pressure and economy of radial engines should 
not be as good as those of any other type. 

A B C. — The development of air-cooled fixed radial engines 
has been carried on in England more than elsewhere. The 




Fig. 146. — Sectional outlines of A B C "Dragonfly." Radial engine. 

ABC engines, built by the Walton Motors Co., have the following 
general characteristics : 



Type name 



Gnat 



Wasp 



Dragonfly 



Number of cylinders (copper-coated 

steel fins) 

Bore, inches 

Stroke, inches 

Normal brake horse power 

Revolutions per minute 

Oil consumption, pints per hour . . 
Gasoline per brake horse power 

hour, pints 

Weight of engine, dry, pounds 

Weight per brake horse power, 

pounds 

Over-all diameter, inches 



2 

4.75 
5.5 
45 
1,800 
1.7 

0.56 
115 

2.3 



7 

4.75 
6.25 
200 
1,800 
4 

0.56 
320 

1.6 

42.7 



9 

5.5 
6.5 
340 
1,650 
7 

0.56 
600 

1.765 
50.5 



RADIAL AND ROTARY ENGINES 



195 



The Wasp and Dragonfly engines have each two exhaust valves 
and one inlet valve per cylinder. Their engines use the master- 
rod connecting-rod assembly with roller bearings (see p. 207) and 
have counterbalance weights. Sectional views of the Dragonfly 
engine are shown in Fig. 146. 

Cosmos. — The Cosmos Engineering Co. has fixed radial engines 
with the following characteristics : 



Type name 



Lucifer 



Jupiter, 
direct 
drive 



Jupiter, 
geared 



Mercury 



Hercules, 
geared 



Number of cylinders . . . 

Number of rows 

Bore, inches 

Stroke, inches 

Normal brake horse 



power 

Brake mean effective 
pressure, pounds per 

square inch 

Revolutions per minute . 
Propeller speed, revolu- 

tins per minute 

Weight of engine, dry, 

pounds 

Weight per brake horse 

power, pounds 

Weight per brake horse 

power at maximum 

power, pounds 

Over- all diameter, inch. 



3 
1 

5.75 
6.25 



9 
1 

5.75 
7.5 



100 



1,600 



400 



113 

1,650 



220 

2.2 



636 
1.59 



1.413 
52.5 



9 
1 

5.75 
7.5 



450 

113 
1,850 

1,200 

757 



52.5 



14 
2 
4.375 

315 



1,800 

587 
1.863 

41.625 



18 
2 

6.25 
7.5 



1,000 

1,750 

1,150 

1,400 

1.4 



The Jupiter engine has two exhaust and two inlet valves; the 
Mercury engine has two exhaust and one inlet valve. 

Performance curves for the Jupiter engine are shown in Fig. 
147. It will be seen that the brake mean effective pressure 
reaches a maximum of 117 lb. per square inch at 1,700 r.p.m. 

A special feature of the Jupiter engine is the method of con- 
veying the explosive charge to the cylinders. There are three 
independent carburetors at the rear of the engine discharging 
into the cover of the annular inlet chamber which forms the rear 
of the crankcase. This chamber (Fig. 148) contains an alu- 
minum spiral casting which fits closely into the chamber. The 



196 



THE AIRPLANE ENGINE 



580 
560 
540 
520 
500 
480 
460 
440 
420 
.•400 

X380 

cti 

360 

340 
320 
300 
280 
260 
240 
220 
200 



.. — ,h — *■ , 



120 

Q." 

110^ 



"1000 1200 1400 1600 1800 2000 

R.P. M. 

Fig. 147. — Performance curves of Cosmos "Jupiter' 



2200 

radial engine. 




Fig. 148. — Induction chamber of Cosmos "Jupiter." 



RADIAL AND ROTARY ENGINES 



197 



casting constitutes a three-part spiral. The carburetors dis- 
charge into the spaces marked X, Y and Z respectively. The 



Cenfrifuga/'V 
pump 



Carburetor 




Fig. 149. — Longitudinal section of Salmson radial engine. 



space X is part of the spiral marked AAA, so that the mixture 
drawn into X will flow along the spiral groove AAA. This 



198 



THE AIRPLANE ENGINE 



groove is opposite the inlet pipes for cylinders 2, 8 and 5; similarly 
the middle carburetor will supply cylinders 3, 9 and 6. ' This 
arrangement gives the mixture a clean sweep from the carburetor 
to the cylinder and isolates the cylinders in three groups so that 
should one carburetor fail to act properly there would still be six 
cylinders in normal action. 

Ex ha 'us f Valve--\ v' In let Valve 




Fig. 150. — Transverse view of Salmson radial engine. 

Salmson. — The Salmson (Canton-Unne) engine is a good 
example of the water-cooled fixed-radial engine. Figure 149 shows 
a longitudinal section of a nine-cylinder engine; Fig. 150 is a 
transverse view of the same engine. The general dimensions 
of the engine are: bore, 125 mm.; stroke, 170 mm.; ratio of 
compression, 5.3; weight of engine without water or radiator, 



RADIAL AND ROTARY ENGINES 



199 



474 lb.; weight of water in jackets 20 lb.; power at 1,500 r.p.m., 
250 h.p.; weight dry per horse power, 1.89 lb.; gasoline con- 
sumption per horse-power hour, 0.507 lb.; oil consumption per 
horse-power hour, 0.077 lb. The variation of brake horse power 
with engine speed is shown in Fig. 151. 

The cylinders are steel forgings 3 mm. thick; the jackets 
are of sheet steel welded to the cylinders. The inlet and exhaust 
valves are symmetrically located and are both 62.5 mm. dia- 
meter; they are held to their seats by rat-trap springs. The 
connecting-rod assembly is of the master-rod type (see p. 203) 
with ball bearings on the crankpin. The crankshaft is of the 



300 



^275 



£250 

CD 



225 



1300 



1400 1500 

Revolutions per Minute 



1600 



1700 



Fig. 151. — Performance curve of Salmson radial engine. 



built-up type with counterweights. The valves are operated 
through push rods and rocker arms from a cam sleeve which is 
revolved on the crankshaft at one-fourth the engine speed by 
means of an epicyclic gear set. There are three pairs of cams on 
the cam sleeve, each pair at opposite diameters in its own plane. 
In each of the three planes are the cam followers of both valves 
for three cylinders. All the valves will be opened twice in one 
revolution of the cam sleeve. Consequently in two revolutions 
of the engine, or one-half revolution of the cam sleeve, each of the 
valves will have been operated once. 

The inlet valve opens at top center and closes 55 deg. after the 
bottom center, or at about 16 per cent of the return stroke. 
Ignition occurs about 30 deg. before top center. The exhaust 
opens 65 deg. before bottom center and closes at top center. 



200 



THE AIRPLANE ENGINE 



The water circulation is shown in Fig. 152. A centrifugal 
pump taking water from the bottom of the radiator discharges 
it through two pipes into the heads of the two lowest cylinders. 
The top and bottom of each jacket is connected by pipes to the 
tops and bottoms respectively of the adjacent cylinders. The 
water is finally delivered from the top of the highest cylinder to 
the radiator. 

The carburetor (Zenith) discharges through long vertical pipes 
into the annular inlet chamber at the rear of the crankcase and 
thence through separate inlet pipes to the individual cylinders. 



Wafer Screen—., 



Thermometer 



Radia-tor 





Fig. 152. — Water circulation in Salmson radial engine. 



The exhaust passes from each cylinder into a sheet metal exhaust 
duct which encircles the engine, discharges at the sides of the 
fuselage, and is stream-lined to serve as a cowling for the engine. 
Details of Radial and Rotary Engines. — Air-cooled cylinders 
are either made from solid steel, as in the Gnome, Le Rhone and 
Clerget rotary engines, or they are composite with steel barrel 
or liner and aluminum alloy head. All-aluminum cylinders have 
been tried with fair success but there is doubt of their durability; 
they are no lighter than the other types and their considerable 
longitudinal expansion increases tappet clearances and alters 
valve timing to a greater extent than with other constructions. 



RADIAL AND ROTARY ENGINES 



201 



The satisfactory operation of an air-cooled cylinder depends 
on keeping down its temperature. When overhead valves are 
used, this temperature is highest in the middle of the head. 
With open exhaust as in rotary engines and in some radial engines 
there is not much difficulty in arranging for adequate cooling of 
all parts of the cylinder. 

With overhead valves it is essential 
to make the cylinder head of the 
best available conductor (see p. 346), 
which in practice turns out to be an 
aluminum-copper alloy. The valve 
seats and the working surface of the 
cylinder barrel must be of some harder 
material. When an aluminum head 
is used the valve seats should consist 
of rings of steel or bronze, cast or 
expanded into position. Bronze seats, 
in consequence of their high coefficient 
of expansion, are less likely to come 
loose than steel seats. 

One type of construction is shown 
in Fig. 153. An aluminum casting 
forms the head and surrounds the 
greater part of the steel liner, which 
is shrunk into the casting at about 
300°C. Cylinders of this type have 
given excellent results, but the differ- 
ence between the coefficients of ex- 
pansion of the Steel and aluminum Fig. 153.— Aluminum air-cooled 
. cylinder with steel liner. 

tends to cause separation of the lmer 

and casing at working temperatures and a film of oil may work in 
between them. With cylinders below 4 in. in diameter there 
is little trouble. A shrinkage allowance of about 1 in 600 
should be made. The expansion trouble can be overcome by the 
use of bronze liners if a bronze sufficiently hard to resist wear 
is developed. The holding-down bolts go through lugs in the 
aluminum casting. 

Screwed-in liners have not given good results owing to the 
impossibility of maintaining adequate contact between liner 
and casing. If the contact is good when cold, the difference of 
expansion when hot causes contact at points only. 




202 



THE AIRPLANE ENGINE 



The best method of composite construction is one with an 
aluminum cylinder head into which is cast or screwed a steel 
barrel with its own cooling fins (Fig. 154). This construction is 
mechanically sound and has been used successfully with cast-in 
barrels for sizes up to 6 in. in diameter and with screwed-in barrels 
up to 5^ in. in diameter. The length of the screwed portion 
should be about one-fourth of the cylinder diameter. The 

holding-down bolts grip a ring integral 
with the steel barrel and thereby avoid 
the breakages of holding-down lugs 
which have been rather frequent with 
the construction of Fig. 153. In an- 
other type of construction the barrel 
and head are formed of steel in one 
piece and an aluminim cap embody- 
ing the inlet and exhaust ports is 
bolted to the cylinder head. 

Tests of all-steel cylinders such as 
are used in Le Rhone and Clerget 
engines, with cylinder diameters rang- 
ing from 4 to 6 in., show that the all- 
steel cylinder gives very appreciably 
higher fuel consumption and lower 
brake mean effective power than 
does the aluminum-headed cylinder. 1 
A 53^ by 6^-in. steel cylinder with 
one aluminum inlet and two cast-iron 
exhaust ports bolted to it was 
changed (1) by having an aluminum cap bolted to its head and 
(2) by having the original head cut off and an aluminum 
head cast on to the same barrel. Tests showed that under 
maximum load conditions at 1,450 r.p.m. and in a wind of 82 
miles per hour the aluminum headed cylinder gave 15 per cent 
more power than either of the others. The fuel consumption 
was 26 per cent less than that of the steel cylinder and 20 per cent 
less than that of the capped cylinder. 

A capped steel cylinder is usually not much better than the 
normal steel cylinder; however well fitted initially, " growth" 
and distortion of the aluminum impair the contact after a few 
hours' running. 

1 A. H. Gibson, Inst. Aut. Eng., Feb., 1920. 




Fig. 154. — Steel air-cooled cyl 
inder with aluminum head. 



RADIAL AND ROTARY ENGINES 203 

The largest all-steel air-cooled cylinder tested by Gibson was 
6 by 8 in. With a compression ratio of 4.48 and in a wind of 
75 miles per hour this cylinder developed 115 lb. brake mean 
effective pressure on a fuel consumption of 0.68 lb. per brake- 
horse-power hour at 1,250 r.p.m., and 105 lb. brake mean effective 
pressure at 1,600 r.p.m. An aluminum-headed cylinder of the 
same dimensions developed under the same conditions 121 lb. 
brake mean effective pressure on a consumption of 0.56 lb. per 
brake horse power per hour. 

Cylinder distortion may arise from the fact that the cooling 
air blast is directed against one side of the cylinder. Such 
distortion is negligible when the blast is directed on the exhaust 
side. This side is normally the hottest and needs most cooling. 
Tests on a 5^-in. aluminum cylinder with the blast on the 
exhaust side showed a maximum temperature difference between 
the front and back of the barrel of 58°C, and a mean difference 
of 19°C. With the blast on the inlet side the maximum tem- 
perature difference was 180°C. and the mean 120°C. In spite of 
this the cylinder, which was fitted with an aluminum piston of 
only 0.025-in. clearance, gave no sign of binding, showing that 
even in this extreme case the distortion was not serious. 

With longitudinal fins and a comparatively uniform distribu- 
tion of air flow, the distortion is not noticeably less. The 
exhaust side will be the hottest and the temperature will be less 
uniform than with circumferential fins and a free blast on the 
exhaust side. Furthermore, longitudinal fins do not stiffen the 
cylinder as strongly against distortion as do circumferential fins. 

Connecting-rod Assembly. — The problem of connecting seven 
or nine big-ends to a single crankpin is usually solved either by 
the "articulated or master rod" assembly or by the " slipper" 
assembly. 

The master rod assembly is used on the Gnome and Clerget 
rotaries and on most of the radials. Details of the assembly, as 
installed in the Salmson engine, are given in Figs. 155 and 156. 
The big end of the master rod encircles the crankpin, holds the 
wristpins for all the short rods, and carries the outer races of the 
ball bearings. It will be seen that this construction shortens the 
effective length of all rods except the master rod; that the axes 
of the short rods pass through the crankpin only twice in the 
revolution; and that the obliquity of the short rods is considerably 
greater than that of the master rod. 



204 



THE AIRPLANE ENGINE 



The slipper type of assembly is used in the Le Rhone and 
Anzani engines. The crankpin carries on ball bearings (Figs. 




Fig. 155. — Articulated connecting-rod assembly. 

157 and 158) two thrust blocks each of which has three annular 
grooves lined with bearing metal. The two discs are fastened 




Fig. 156. — Section through articulated connecting-rod assembly. 

together with the annular grooves opposite one another. The 
big ends of the nine connecting rods are provided with slippers 



RADIAL AND ROTARY ENGINES 



205 




Fig. 157. — Section through slipper Fig. 158. — Assembly of slipper type connect- 
type connecting-rod assembly. ing rods. 




Fig. 159. — Diagram of rotary engine with slipper" type connecting-rod assembly. 



206 



THE AIRPLANE ENGINE 




Fig. 160.- 



-Projected views of slip- 
pers. 



each of which is turned with the same radius of curvature as one 
of the annular grooves. Three connecting rods act on each 
groove and consequently there are three designs of slipper. The 
slippers for the middle and outermost grooves are slotted to 
avoid contact with the connecting rods for the innermost and 
middle grooves. The arrangement is shown in outline in Fig. 
159. The plan of the slippers in Fig. 160 shows the slotting to 
prevent interference with adjacent connecting rods. 

The slipper assembly is considerably heavier than the master- 
rod type and consequently is better adapted to rotaries than to 
radials. It has the advantage that the connecting rod is of 

maximum length and conse- 
quently of minimum angularity 
and also that the thrust (or ten- 
sion) of the rod always passes 
through the center of the crank- 
pin. Furthermore a large bear- 
ing surface is provided at the 
thrust block which is easily 
lubricated by the oil thrown off 
from the ball bearings. 
Dynamical Comparison of Radial and Rotary Engines. — 
The fixed-radial engine presents the special problem of a large 
mass rotating with the crankpin and consequently large centrif- 
ugal force. The inertia forces of the reciprocating parts are 
additive to this. The result is a considerable total pressure 
on the crankpin, which is relieved somewhat by the gas pres- 
sures during the explosion strokes. Roller or ball bearings are 
necessary at the crankpin if high speeds of rotation are to be 
maintained. 

Balancing of the primary inertia forces of a single-crank fixed- 
radial engine is readily effected by a mass, approximately equal to 
half the mass of all the reciprocating parts, used as a counter- 
balance opposite the crankpin at crankpin radius. The counter- 
balance weight will add 7 to 10 per cent to the weight of the 
engine and can be avoided only by using two rows of cylinders 
and a double-throw crank. In the last case there is an unbal- 
anced primary couple. Balancing the centrifugal and inertia 
pressures on the crankpin has been accomplished in an ingenious 
manner in the latest design of Cosmos " Jupiter" engine. Two 
bob-weights are suspended on the outer sides of the master rod; 



RADIAL AND ROTARY ENGINES 



207 



their other ends are connected to the main crankshaft balance 
weights through hardened blocks working in slots machined in 
the bob-weights. The bob- weights serve not only to relieve the 
pressure on the crankpin but also as part of the weight necessary 
to balance the engine as a whole. The general arrangement of 
these bob- weights is shown in Fig. 161. 

In rotary engines the pistons and connecting rods rotate about 
a stationary crankpin with an angular velocity which is variable. 
With the master-rod type of connecting-rod assembly there is 




Fig. 161. — Balanced connecting rod of Cosmos "Jupiter" engine. 



some lack of centrifugal balance at the crankpin, but it is usually 
negligible. The connecting rods are subjected to centrifugal 
tensional loading; the pressures on the pins at the ends of the 
rods increase as the square of the revolutions per minute. With 
the same connecting rod loading a fixed-radial engine may run at 
approximately twice the speed of a rotary engine with the same 
moving parts; before that speed is reached, however, the crankpin 
loading of the radial becomes excessive. 

Ball and roller bearings for crankpins of radial engines offer 
special problems. The bearing rotates as a whole and presents 



208 THE AIRPLANE ENGINE 

conditions of loading quite unlike those of stationary bearings. 
Considerable investigation of this matter was made by the 
British Department of Aircraft Production 1 as a result of the 
failure of both caged and uncaged bearings. 

An analysis of the situation showed that with cageless bearings 
the balls are crowded away from the center of rotation, by centrif- 
ugal force, and rub against one another. The balls rotate 
usually at about 2,500 r.p.m. about their own centers; the points 
that touch are always moving in opposite directions and the 
abrasion is considerable. When a cage is used the centrifugal 
force on the cage and the balls causes a displacement of the cage 
until the bearing load On the balls nearest the crank center due to 
the cage wedging between them is equal to the total centrifugal 
load on the cage. This causes a heavy abrasive action between 
the balls and the cage. 

For successful operation it is necessary to have a cage which 
will carry independently the rubbing loads on each ball due to 
centrifugal force. To accomplish this (1) the cage must be 
strong enough to take the independent loads from the balls 
without distortion, (2) sufficient bearing surface must be provided 
at the surface of location of the cage to carry safely the total 
centrifugal load, (3) sufficient bearing surface must be provided 
between the balls and the cage to prevent wear on the cage, (4) 
the cage must be made of a metal of minimum abrasion, and (5) 
all surfaces must run with a continuous flow of oil. 

To meet these requirements a cage as in Fig. 162 may be 
used. This type of cage must be definitely located and not 
displaceable by centrifugal force for more than a few thousandths 
of an inch. The design shown is made in two halves with eight 
hemispherical holes with 0.10 in. clearance for the balls. The 
outside circumference is turned in a flat V to avoid the actual ball 
path, and on either side of the V is a true cylindrical surface 
about % 6 m - wide. The cage is of phosphor bronze and fits 
the outer ball race with a clearance of 0.005 to 0.007 in. This 
bearing proved entirely satisfactory on the crankpin of a 10- 
cylinder, 115 by 150-mm. radial Anzani engine developing 150 
h.p. at 1,300 r.p.m. 

The details of a satisfactorily located cage for rollers for the 
crankpin for a 320-h.p. nine-cylinder radial engine making 1,700 

1 J. B. Swan, The Automobile Engineer, July, 1919. 



RADIAL AND ROTARY ENGINES 



209 



r.p.m. are given in Fig. 163. The cage is located on the roller 
track, which in practice works out advantageously in polishing 
the track and keeping it free from foreign matter. 




0. 568 '"Spherical md 
on 3 3 /i6"diam.PC 



Sec-Hon A- A -A 



View of R.H.HctfF 
Fig. 162. — Located cage for ball bearings. 



Details of successful and unsuccessful ball and roller bearings 
are given in the following table. At speeds above 1,600 r.p.m. 
and with a radius of rotation above 2J-^ in. cageless bear- 



u 0.9_8'' 

\h±ad®\ 



W-rO.OlO 









* £ 




Fig. 163. — Located cage for roller bearings. 



ings will not run satisfactorily if the balls or rollers are larger 
than % in. in diameter and the inner race larger than 1.5 in. in 
diameter. 



14 



210 



THE AIRPLANE ENGINE 





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CN CO 00 


Roller or 
ball diam- 
eter in 
inches 


3 s "g 

H O ^ 
® ^ \» 
V \<)" i-N 


Revolu- 
tions per 
minute 


O O O 

o o o 

00 t^ CO 


CD CO 

O rd 

u o 

ui .9 


o o o 

OS >o H 
iO CO CO 


. CO 
CD <» 

k •# 

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W d 


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»o »o »o 

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CO 0J 


o o o 

t^ CN »0 

HfOH 


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d 
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d 
<u 

"o 

01 

ft 
>> 
H 


^a 

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s 

c 
1 


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i 



RADIAL AND ROTARY ENGINES 211 

The crankshaft in a radial engine will be solid or built-up 
according as the big end of the connecting rod has a plain bearing 
or a ball or roller bearing. The plain bearing is likely to give 
trouble in view of the heavy loading of the bearing except for 
low-speed engines, and can be used only with high-pressure 
forced lubrication; ball or roller bearings are very generally used. 

The built-up crank necessary with ball or roller bearings may 
be either a two-web shaft with equal loading on front and rear 
bearings or an overhung crank with a drag crank for driving 
auxiliaries. The former type is the more desirable. With the 
overhung crank the main balance weight is on a single-crank web, 
which leads to an unbalanced couple, and also, the diameter of 
shaft and bearings has to be made greater. 

Valve Operation. — Apart from the design of multilobed cams 
and their driving gears, the valve operation of radial engines does 
not present any special problems. In rotary engines the effects 
of centrifugal force on the push rods and tappets have to be met. 
Counterweights have sometimes been used on the valve side of 
the rocker arm, but their use increases the load and wear on the 
cam profile. The necessity for keeping down the over-all engine 
diameter is likely to result in the selection of an unfavorable type 
of valve spring and an undesirable reduction in the length of the 
valve stem guide; failures have been frequent when volute valve 
springs have been used. 

Lubrication. — Rotary engines are always wasteful of oil, 
using almost Jio 1D - P er brake-horse-power hour. There is no 
return of surplus oil to the pump, which consequently has to 
determine the amount of oil used. Plunger pumps are always 
used discharging directly to the main bearings, big end and cam 
gear, and relying largely on centrifugal force for the lubrication 
of wristpins and cylinders. 

In radial engines either a plunger pump or a gear pump may 
be used with a dry sump. There is danger of over-oiling the 
lower cylinders. Most of the oil goes direct to the crankpin and 
is distributed thence by centrifugal force to the bearings, connect- 
ing-rod assembly, and cylinders. The oil consumption of radial 
engines runs from about 0.02 to 0.04 lb. per brake -horse-power 
hour. 



CHAPTER IX 
FUELS AND EXPLOSIVE MIXTURES 

The properties desired in an airplane engine fuel are as follows: 

1. It must have a high heat of combustion per pound. This 
determines the cruising radius for a given weight of fuel, since 
efficiency does not vary appreciably with the fuel. 

2. It must have a high heat of combustion per cubic foot of 
explosive mixture if it is to develop high horse power per cubic 
foot of piston displacement. Alcohol and gasoline have about 
the same heats of combustion per cubic foot of explosive mixture 
but very different heats of combustion per pound of fuel. The 
heat of combustion is nearly constant for all the available fuels. 

3. It must be able to withstand high compression without 
preignition or detonation. 

4. It must vaporize readily (preferably with little or no pre- 
heating of the air) upon admixture with air and should be com- 
pletely vaporized at the beginning of explosion. For good 
distribution it should be completely vaporized upon reaching 
the admission manifold, but this is not usually attained. 

5. Combustion should be complete, leaving no solid residue in 
the cylinder. 

6. The fuel and products of combustion must not be corrosive. 

7. The explosion rate must be neither too rapid (as with 
hydrogen, acetylene and ether) nor too slow. 

8. The bulk of fuel and the weight of the container must be 
low. This eliminates gaseous fuels. 

The liquid fuels which meet the above conditions best are: 
(a) certain hydrocarbons, which form the constituents of gasoline 
and of certain coal tar products, and (6) the alcohols. The 
hydrocarbons under consideration may be divided into two main 
groups, saturated and unsaturated. The latter term is here 
applied to the behavior and not to the composition of the sub- 
stance. The saturated hydrocarbons are again subdivided into 
the aliphatic or acyclic group, and into the aromatic or cyclic 
group. The hydrocarbons all form series in which the members 
differ from each other by the addition of CH 2 . The members of 
the groups of most importance are listed in Table 9, together with 

212 



FUELS AND EXPLOSIVE MIXTURES 



213 



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214 THE AIRPLANE ENGINE 

the boiling points (temperatures of vaporization at atmospheric 
pressure), densities as compared with that of water and latent 
heats of evaporation. 

Certain properties of these compounds are given in the follow- 
ing tables, in which there are also included, for convenience, the 
properties of air, and its constituents, and of the products of 
combustion of the fuel elements. In Table 10 data are given for 
the gaseous state only. Column 3 gives the density of each 
substance compared with air at the same pressure and tempera- 
ture; column 4 gives the weight in pounds per cubic foot of the 
substance. Some of the quantities there given are fictitious in 
that the substance is not gaseous at 32°F. and 760 mm., but the 
quantity is of value in permitting the easy determination of 
specific weight at those temperatures and pressures at which it is 
gaseous. Column 5 is the reciprocal of column 4. The quantity 
R is the constant in the gas equation pv = wRT. 

In Table 11 are given combustion data for the fuel constituents. 
Column 4 gives the volume of air necessary to burn one volume of 
the gaseous fuel, both being at the same pressure and tempera- 
ture; the products of complete combustion are in all cases CO2, 
H 2 and N 2 and their volumes are given in columns 5, 6 and 7. 
The mixture usually experiences a change in volume as a result 
of the chemical changes resulting from combustion, entirely 
independent of the change in pressure and temperature; this 
change in volume is given in column 8. Columns 9 to 12 give the 
weight of air required for combustion of 1 lb. of fuel and the 
weights of each of the resulting products. 

Table 12 gives the heat of combustion of each of the substances 
listed. There is also given the heat of combustion per pound 
of explosive mixture, and the heat of combustion per cubic foot 
of explosive mixture measured at 60°F. and standard atmospheric 
pressure, the mixture being assumed to contain only that amount 
of air which is chemically necessary. It will be noted that two 
values of heat of combustion are given under each head, a higher 
and a lower heat value, and that these are different for all the 
fuels which contain hydrogen. The higher heat value is the 
total heat liberated by combustion, or the heat which would 
be given up by the mixture when burned in a closed vessel 
and cooled to its initial temperature. Whenever there is 
hydrogen in the fuel, water is formed by combustion and part 
of the heat of combustion is absorbed in vaporizing it. If, 



FUELS AND EXPLOSIVE MIXTURES 



215 



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216 



THE AIRPLANE ENGINE 



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FUELS AND EXPLOSIVE MIXTURES 



217 



Table 12. — Heats of Combustion 



Substance 



Chemical 
formula 



B.t.u. per pound 



High Low High 



B.t.u. per cubic 
foot of theoreti- 
cal mixture at 
60°F. and 760 
mm. 



Low 



B.t.u. per 
pound of theo- 
retical mixture 



High Low 



Hydrogen 

Carbon-monoxide 

Methane 

Ethane 

Propane 

Butane 

Pentane 

Hexane 

Heptane 

Octane 

Nonane 

Benzene 

Toluene 

Cyclo-hexane. 

Ethylene 

Propylene 

Butylene 

Acetylene 

Allylene 

Naphthalene. . . . 
Methyl alcohol . . . 
Ethyl alcohol. . . . 



H 2 

CO 

CH 4 

C 2 H 6 

CaH 8 

C4H10 

C5H12 

CeHi4 

C7H16 

CsHis 

C9H20 

CeHe 

C 7 H 8 

C 6 Hi 2 

C2H4 

C3H6 

C4H8 

C2H2 

C3H4 

CioH 8 

CH4O 

C 2 H 6 



62,100 


52,920 


95.6 


4,380 


4,380 


94.6 


23,850 


21,670 


95.8 


22,230 


20,500 


99.7 


21,600 


20,055 


101.2 


21,240 


19,780 


101.7 


21,140 


19,600 


103.0 


20,800 


19,380 


102.5 


20,600 


19,200 


102.0 


20,400 


19,020 


101.4 


20,380 


19,015 


101.7 


18,070 


17,400 


101.0 


18,250 


17,490 
18,900 


100.8 


21,600 


20,420 


104.9 


21,330 


20,150 


104.8 


20,880 


19,700 


104.4 


21,600 


21,020 


115.0 


21,200 


20,325 


112.1 


17,410 


16,860 


101.1 


9,550 


8,460 


99.5 


13,000 


11,650 


105.2 



81.5 
94.6 
87.0 
91.8 
94.0 
94.6 
95.6 



4 

5 

2 
5 
5 
2 
3 

98.5 
112.0 
107.0 
98.0 
88.2 
94.3 



1,760 
1,265 
1,310 
1,300 
1,297 
1,285 
1,297 
1,284 
1,275 
1,266 
1,266 
1,184 
1,261 

1,371 
1,356 
1,325 
1,515 
1,434 
1,250 
1,281 
1,303 



1,500 
1,265 
1,190 
1,200 
1,205 
1,203 
1,203 
1,197 
1,188 
1,176 
1,184 
1,140 
1,208 

1,295 
1,280 
1,254 
1,475 
1,375 
1,210 
1,133 
1,167 



as is usual in gas engines, the gases escape at so high a tem- 
perature that no water vapor is condensed in the cylinder, 
the latent heat of vaporization of the water is not available for 
raising the temperature of the products of combustion, or for 
doing work. The useful heat of combustion is consequently the 
total heat less the heat absorbed in vaporizing the H 2 formed 
by the combustion. The heat of vaporization depends on a 
number of factors. A value of 950 B.t.u. per pound may be 
assumed and if this is multiplied by the number of pounds of 
water formed per pound of fuel burned it will give (with an 
approximation adequate for ordinary purposes) the unavailable 
heat. The lower heat value is the total heat minus the unavail- 
able heat and is the value commonly used by engineers in dealing 
with gas engine problems. 

Gasoline. — The gasolines at present on the market are of three 
different types: 1 



1 Dean, Motor Gasoline, Technical Paper 166, U. S. Bureau of Mines. 



218 THE AIRPLANE ENGINE 

1. "Straight" refinery gasoline. 

2. Blended casing-head gasoline. 

3. Cracked and blended gasoline. 

"Straight" refinery gasoline is produced by distillation. 
Crude petroleum is first distilled from a fire still, and the con- 
densed product is collected until it reaches some predetermined 
density. This so-called crude naphtha or benzine is then acid- 
refined and steam-distilled. Several products of different ranges 
of volatility may be produced or the steam distillation may simply 
separate the product from the less volatile " bottoms.' ' Straight 
refinery gasolines consist mainly of aliphatic hydrocarbons (see 
Table 9) and are generally characterized by a low content of 
unsaturated and aromatic hydrocarbons and by a distillation 
range free from marked irregularities. 

Blended casing-head gasolines are of recent development. 
Casing-head gasoline is obtained, by compression or absorption, 
from natural gas and is too volatile for general use. Before 
marketing, it is generally blended with enough heavy naphtha to 
produce a mixture that can be handled safely. As a result of this 
blending, the volatility range is usually characterized by a con- 
siderable percentage of constituents of both low and high boiling 
points and a lack of intermediate constituents. Skilful blending 
may change this characteristic. 

In chemical properties the blended casing-head gasoline seems 
to be identical with a straight refinery product of the same 
distillation range. 

Cracked or synthetic gasoline is also a recent development. 
An oil consisting mainly of heavier hydrocarbons is subjected to 
high temperature and pressure and is thereby broken down or 
"cracked" into lighter constituents. This cracked gasoline 
is generally marketed in the form of blends with refinery and 
casing-head gasoline. Cracked gasolines differ chemically from 
straight-refinery and casing-head gasolines in having a larger 
amount of unsaturated and aromatic hydrocarbons. The heat 
of combustion of the aromatic compounds averages about 15 per 
cent less than that of the acyclic compounds, but, as shown on 
page 238, the presence in moderate amounts of certain aromatic 
compounds may improve the thermal efficiency of the engine 
enough to offset any disadvantage, for airplane use, of a lower 
heat of combustion. 



FUELS AND EXPLOSIVE MIXTRES 



219 



Specifications. — In the past, gasolines have usually been 
described and bought on a gravity specification. So long as the 
gasolines in the market were straight refinery products such a 
specification was reasonably satisfactory, but, with the develop- 
ment of blended casing-head gasolines, it has become impossible 
to determine the volatility of a gasoline by density measurements. 
A given density may represent either a narrow "cut" consist- 
ing of a product which evaporates with a very narrow range of 
temperature, or a mixture of a volatile low-density product with a 
product of high density and low volatility. The former fuel 
might be admirable for airplanes while the latter might be quiet 
unsuitable. 

Specific gravity is best expressed as the ratio of the density 
of the fuel to that of water, both at 60°F. The trade practice 
has been to use the Baume scale of density. This arbitrary 
scale has nothing to recommend it, and suffers the disadvantage 



60° 
Table 13. — Specific Gravities at w^s F. Corresponding to Degrees 







Baume for Liquids Lighter 


than Water 






1 

3 

d 

pq 

m 

O! 
<D 

S> 

P 


1 
g 

© 

1 
P. 

CO 


a 

3 
03 

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m 

<a 

<D 
ft 

p 


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g 

a 
«d 
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ft 

CO 


a 

PQ 

OS 

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&> 

4) 

P 


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*> 

03 
u 

to 
o 
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ft 

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a 

oJ 

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00 

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a 

03 

pq 

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a> 

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U 
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a> 

P 


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03 
& 

•3 

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ft 

CO 


25 


0.9032 


40 


0.8235 


55 


0.7568 


70 


0.7000 


85 


0.6512 


26 


0.8974 


41 


0.8187 


56 


0.7527 


71 


0.6965 


86 


0.6482 


27 


0.8917 


42 


0.8140 


57 


0.7487 


72 


0.6931 


87 


0.6452 


28 


0.8861 


43 


0.8092 


58 


0.7447 


73 


0.6897 


88 


0.6422 


29 


0.8805 


44 


0.8046 


59 


0.7407 


74 


0.6863 


89 


0.6393 


30 


0.8750 


45 


0.8000 


60 


0.7368 


75 


0.6829 


90 


0.6364 


31 


0.8696 


46 


0.7955 


61 


0.7330 


76 


0.6796 


91 


0.6335 


32 


0.8642 


47 


0.7910 


62 


0.7292 


77 


0.6763 


92 


0.6306 


33 


0.8589 


48 


0.7865 


63 


0.7254 


78 


0.6731 


93 


0.6278 


34 


0.8537 


49 


0.7821 


64 


0.7216 


79 


0.6699 


94 


0.6250 


35 


0.8485 


50 


0.7778 


65 


0.7179 


80 


0.6667 


95 


0.6222 


36 


0.8434 


51 


0.7735 


66 


0.7143 


81 


0.6635 


96 


0.6195 


37 


0.8383 


52 


0.7692 


67 


0.7107 


82 


0.6604 


97 


0.6167 


38 


0.8333 


53 


0.7650 


68 


0.7071 


83 


0.6573 


98 


0.6140 


39 


0.8284 


54 


0.7609 


69 


0.7035 


84 


0.6542 


99 
100 


0.6114 
0.6087 



220 



THE AIRPLANE ENGINE 



that the greater the density the lower is the number of "degrees" 
on the Baume scale. For liquids lighter than water, the relation 
between the specific gravity and Baume scale, B, is given by 
the expression 

Specific gravity = 13Q , B 

Numerical values are given in Table 13. Commercial gasolines 
range from about 55 to 75°Be. (Sp. gr. 0.758 to 0.684). The 
Eastern gasolines are lightest (60 to 75°Be.) and the California 
gasolines heaviest (57 to 63°Be.). 



Cork 

X Thermometer 

<./' Cooling Trough 

Cork- / Containing Cracked Ice and Wafgr 




Fig. 164. — Distillation apparatus. 

Volatility. — Volatility is the basic property that determines the 
grade and usefulness of a gasoline. The presence of low-boiling 
constituents is desirable to permit easy starting of a cold motor 
but may result in excessive evaporation losses in the commercial 
handling of the fuel. 

The volatility is determined by distillation (Fig. 164). A 
100-gram sample of the fuel is heated slowly while the vapor given 
off is condensed and collected. The first drop of gasoline should 
fall from the end of the condenser tube in 5 to 10 min. The 



FUELS AND EXPLOSIVE MIXTURES 



221 



rate of evaporation is kept about 4 c.c. per minute. A 

thermometer indicates the temperature of the vapor above the 

fuel. Readings of 

this temperature are 

taken when the first 

drop of distillate falls 

(initial point) and as 

each 10 per cent or 

other selected p e r- 

centage has distilled, 

until at the end a 

dry point is reached. 

The observations are 

usually plotted as in 

Fig. 165 and give a 

record of the volatility 

of the fuel. 

Specifications for 
aviation gasoline have 
been prepared by the U. S. Committee on Standardization of 
Petroleum Specifications 1 and are as follows: 

Gasoline Distillation Test Specifications 




100 125 150 

Temperature , Deg.Ceirr. 

Fig. 165. — Distillation curves for straight refinery 
and casing-head gasoline. 





Aviation 


Aviation 


Grade 


gasoline. 


gasoline, 




domestic grade 


fighting grade 


Thermometer reading range when 5 per 






cent is recovered in receiver 


122 to 167°F. 


122 to 149°F. 


Thermometer reading when 50 per cent is 






recovered in receiver, not more than .... 


221°F. 


203°F. 


Thermometer reading when 90 per cent is 






recovered in receiver, not more than 


311°F. 


257°F. 


Thermometer reading when 96 per cent is 






recovered in receiver, not more than .... 


347°F. 


302°F. 


End-point shall not be higher than 


374°F. 


329°F. 


Distillate recovered in the receiver from 






the distillation at least 


96 per cent 


96 per cent 


When the residue is cooled and added to 






the distillate in the receiver the distilla- 






tion loss shall not exceed 


2 per cent 


2 per cent 







1 Bureau of Mines, Bulletin No. 5, effective Dec. 29, 1920. 



222 THE AIRPLANE ENGINE 

In addition the specifications require for both grades of 
aviation gasoline the following properties: 

Color: Water white. 

Doctor test : Negative. 

Corrosion test: 100 c.c. of the gasoline shall cause no gray or black 
corrosion and no weighable amount of deposit when evaporated in a polished 
copper dish. 

Unsaturated hydrocarbons: maximum proportion of the gasoline soluble 
in concentrated sulphuric acid, 2 per cent. 

Acid heat test: the gasoline shall not increase in temperature more than 
10°F. 

Acidity: the residue after distillation shall not show an acid reaction. 

The gasoline shall be free from undissolved water and suspended matter. 

The Doctor Test is made by shaking two volumes of gasoline 
with one volume of " doctor" solution (sodium plumbite) in a 
test tube, shaking for 15 sec, adding a pinch of flowers of sulphur, 
shaking again and allowing to settle. If the liquid remains 
unchanged in color and the sulphur remains bright or only slightly 
discolored, the test is negative and the gasoline is "sweet." 

The Acid Heat Test is made by adding 30 c.c. of 66° commercial 
sulphuric acid to 150 c.c. of gasoline, both being at room 
temperature. After mixing, shake for 2 min. and observe the 
rise in temperature. 

Volatility curves for three straight refinery gasolines and for 
three blended casing-head gasolines, of approximately the same 
densities, are given in Fig. 165. The casing-head gasolines are seen 
to have larger percentages distilled below 50°C, but have longer 
distillation ranges. This results in a fairly uniform slope of the 
distillation curve. The large percentage unevaporated at 150°C. 
shows that the fuel is a blend or mixture of heavier and lighter 
fuels. 

It should be noted further that the "cut" or fraction distilling 
off at any given temperature will be different from different 
gasolines. This is demonstrated in Fig. 166, which shows that the 
100°C. cut may vary in specific gravity from 0.710 to 0.733 and the 
150°C. cut from 0.748 to 0.780. In other words, volatility alone 
is not sufficient to characterize a gasoline. 

From the volatility curves for straight refinery gasoline, Fig. 
165, it will be seen that the average temperature of evaporation 
(boiling temperature) from a high-grade gasoline is 100°C. From 
the specific gravity curves it appears that the average density at 



FUELS AND EXPLOSIVE MIXTURES 



223 



100°C. is almost 0.7. As the constituents are mainly aliphatic 
hydrocarbons it is safe to assume that a high-grade gasoline con- 
sists principally of hexane (C 6 Hi 4 ) and heptane (C 7 Hi 6 ), whose 
boiling points are 09 and 98.4°C. and densities 0.676 and 0.7, re- 
spectively (see Table 9). 

Calorific Value.— The 
calorific value of com- 
mercial gasoline varies 
very slightly with type 
of fuel, field of origin, or 
density. Exhaustive 
tests by the U. S. Bureau 
of Mines show only 1.5 
per cent difference be- 
tween the highest and 
lowest values for a range 
of density from 0.687 to 
0.745 (73.8 to 57.9°Be\), 
the samples investigated 
including all the com- 
mercial types. The 
average high heat value 
is 20,200 B.t.u. per 
pound. It should be 
noted, however, that 
gasoline is sold by the 
gallon and that there is 
considerable difference 
on that basis; the calorific value ranges from 124,000 B.t.u. per 
gallon for sp. gr. 0.745, to 116,500 B.t.u. for sp. gr. 0.687, a dif- 
ference of over 7 per cent in favor of the heavier fuel. This 
difference is not important in airplane use, since the weight of 
fuel that has to be carried is the important factor, and not its 
volume ; in automobile use it may more than offset the disad- . 
vantages resulting from the use of a less volatile fuel. 

Benzene or benzol (CeH 6 ) is a fuel which has been used con- 
siderably in airplanes, though generally mixed with gasoline. It 
is obtained chiefly from by-product coke-ovens. 

Commercial 90 per cent benzol has a specific gravity of about 
0.88. Its distillation curve should show an initial point not 
lower than 74°C, 90 per cent at or below 86°C, 95 per cent at or 




Fig. 



100 125 150 

Temperature, DegCent 

166. — Density of "cuts" from various gas- 
olines. 



224 



THE AIRPLANE ENGINE 



below 95°C, and end point not above 150°C. With a calorific 
value of 18,000 B.t.u. per pound the heating value per gallon is 
132,000 B.t.u., or considerably higher than that of gasoline. 

When mixed with gasoline there is no change in total volume. 
The distillation curve for such a mixture, containing 20 per cent 
of benzol and 80 per cent high-grade gasoline, is shown in Fig. 
167, together with the distillation curves of the benzol and the 
gasoline. 




First 10 20 
Drop 



90 Dry 
Point 



Percentage Distilled 
Fig. 167. — Distillation curve of benzol-gasoline mixture. 



It will be observed that the effect of the addition of benzol is 
to increase the volatility of the mixture; with 30 to, 50 per cent 
distilled the volatility is greater than that of either of the con- 
stituents, after which it becomes intermediate to the volatility of 
the constituents. This improvement in volatility has been 
found to be distinctly advantageous in increasing engine power 
at high altitudes. The lower heat of combustion of the mixture 
results in the consumption of a greater weight of fuel per brake 
horse-power hour than with straight gasoline. 

Alcohol has been used mixed with gasoline or benzol or both, as 
an airplane fuel. Methyl alcohol, CH 4 (wood alcohol) has a 
boiling point of 65°C. and sp. gr. 0.81, heats of combustion, high 
9,550, low 8,460, B.t.u. Ethyl alcohol, C 2 H 6 (grain alcohol) 
has a boiling point of 78°C, density 0.79, heats of combustion, 
high 13,000, low 11,650 B.t.u. Commercial alcohol, either pure 
or denatured, contains water (e.g., 10 per cent by volume in 90 



FUELS AND EXPLOSIVE MIXTURES 225 

per cent alcohol) and has a higher boiling point than pure alcohol. 
The effect of the addition of alcohol to gasoline is to improve the 
volatility. The calorific value of alcohol is so low compared with 
gasoline that its use inevitably increases the weight of fuel 
burned per unit of power developed. It does not, however, 
diminish the power developed, because the heat of combustion 
per unit volume of explosive mixture (see Table 12) is about the 
same as for gasoline; it may even increase the power output 
slightly. Its use also permits an increase in the permissible 
compression ratio for the engine and thereby improves the 
thermal efficiency. 

Hydrogen has the highest calorific value of any of the fuels, per 
pound, but not per cubic foot of explosive mixture (Table 12). 
Apart from its high cost, it is objectionable on account of the 
great violence of the explosion when mixed with the proper 
amount of air. It cannot be carried in airplanes unless com- 
pressed to very high pressure in steel tanks, which makes the 
fuel system too heavy, or in the liquid form, which increases 
greatly the cost of the fuel. Liquid hydrogen has a temperature 
below — 400°F. and cannot be kept from evaporating rapidly 
without the very best of heat insulation; no sufficiently robust 
container with adequate heat-insulating qualities has been devised 
as yet. Hydrogen gas has been used for starting cold engines. 

It is often necessary to waste some of the hydrogen contained 
in a dirigible balloon. Attempts to burn the hydrogen alone in 
the engine have shown that only about one-third of the maximum 
horse power of the engine could be developed without serious 
detonations. By mixing hydrogen with the gasoline it is possible 
to develop the maximum power without trouble and thereby to 
save gasoline; at the higher powers only a small quantity of 
hydrogen can be burned. 

Acetylene, (C2H2), like Hydrogen, gives explosions of great 

violence in the. cylinder. Its heat of combustion per cubic foot 

of explosive mixture is highest of all the fuels given in 

Table 12. It may be stored either in the gaseous or liquid 

forms, but with the same objections (though to a less degree) 

as hydrogen. It can be generated by adding water to calcium 

carbide, leaving a residue of slaked lime. As the residue is 

considerably heavier than the acetylene, the total weight of the 

fuel system becomes excessive, if it is attempted to generate the 

acetylene in an airplane. 
15 



226 THE AIRPLANE ENGINE 

Ether has, as its principal advantage, the fact that its boiling 
point (35°C.) is lower than that of any of the other available fuels 
which are liquid at ordinary temperatures. This gives it a special 
value in starting a cold engine. Its use has been restricted to 
that purpose. The heat of combustion is rather low. 



EXPLOSIVE MIXTURES 

Properties of Vapors. — Every liquid gives off vapor continu- 
ously until the pressure exerted by that vapor at the surface of the 
liquid reaches a limiting value, which depends, for any given 
liquid, on its temperature only. The vapor liberated is always 
at the temperature of the liquid and it is said to be " saturated" 
when it is at the limiting pressure. The relation between the 
pressure and temperature of a saturated vapor is determinable 
only by experiment. 

The pressure exerted by a vapor will, in time, reach the satura- 
tion pressure if the liquid is contained in a vessel of moderate 
dimensions; the presence, above the liquid, of gases or other 
vapors which are inert to the vapor under consideration and are 
at the same temperature will not affect the saturation pressure. 
The total pressure in the vessel (assuming no change of tempera- 
ture) will be the sum (1) of the pressures of the gases and vapors 
already there and (2) of the saturation pressure of the liquid. 

If the containing vessel is very large or if the time available 
is too short, or if the weight of liquid put into the vessel is less 
than the weight of saturated vapor necessary to fill the vessel, 
the vapor will have a pressure less than the saturated pressure 
and will be in the condition known as " superheated." Suppose 
the temperature of the superheated vapor to be T. If this 
vapor is cooled at constant pressure, with consequent diminu- 
tion in volume, a temperature, T , will eventually be reached 
at which the vapor is saturated. The cooling process is similar 
to that used for determining the dew-point of air; the dew-point 
is the saturation temperature. The vapor is said to be super- 
heated T-T degrees. All unsaturated vapors are superheated. 
When superheated they may be considered to behave like per- 
fect gases. 

A saturated vapor cannot exist, as such, at a pressure greater 
than the saturation pressure. If a saturated vapor is cooled at 



FUELS AND EXPLOSIVE MIXTURES 227 

constant volume, thereby lowering the saturation pressure, 
some of the vapor will condense. If a saturated vapor is com- 
pressed, keeping the temperature constant, condensation will 
take place; if, on the other hand, it is expanded at constant 
temperature it will become unsaturated (superheated) unless 
liquid is present to supply more vapor by evaporation. The 
presence of other inert vapors will not affect these phenomena. 
If air is passed through or over a liquid (as in certain obsolete 
types of carburetor), and if the contact is sufficiently intimate 
and prolonged, the air will leave carrying with it the saturated 
vapor of the liquid. If a liquid is injected into a current of air 
(as in modern carburetors) and if the contact is sufficiently 
intimate and prolonged and if, furthermore, the weight of liquid 
present is sufficient for that purpose, the air will carry its own 
volume of the saturated vapor of the liquid. If more liquid is 
injected than is necessary for this purpose the excess liquid will 
remain in the liquid state. In any case, the total pressure of the 
carbureted air is the sum of the partial pressures of the air and of 
the vapor. If the pressure of the carbureted mixture is p, and 
the pressure of the vapor is p v , and of the air in the carbureted 
mixture is p a , then 

P = Pa + Pv 

The relation between the saturation pressures and temperatures 
of the liquid fuels which are of importance in airplane engines is 
given in Fig. 168. Table 14 gives their values for certain selected 
temperatures. 

The specific volumes (volumes of 1 lb.) of the saturated vapors 
are calculated on the assumption that they are perfect gases. 
This assumption is fairly satisfactory for the low vapor pressures 
which alone are of interest in engine mixtures. Taking, for 
example, heptane (C 7 Hi 6 ) at 60°F., the molecular weight is 7 
X 12 + 16 = 100. The gas constant R is inversely as the mo- 
lecular weight of the gas; taking R for oxygen as 48.25, R for hep- 

32 
tane is j^. X 48.25 = 15.45, and the specific volume at 60°F. 

and at the saturation pressure of 0.54 lb. per square inch is 
RT 15.45 X 520 inQ .. 

V = = n ZA w 1AA = ^ CU. ft. 

p 0.54 X 144 



228 



THE AIRPLANE ENGINE 



The weight of air required for combustion is obtained from the 
chemical equation, 

C 7 Hi 6 + 110 2 = 7C0 2 + 8H 2 
The relative weights of heptane and oxygen are 100 and 11 X 32. 
As the oxygen content of air is 23.4 per cent by weight, the air 

11 X 32 

required for the combustion of 1 lb. of heptane is — ™ — X 

100 

2o~Z — 15.1 lb. The volume of this air at 60°F. and 14.7 lb. 

£0 



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'30 40 50 60 70 80 90 100 110 120 '130 140 150 160 170 180 
Tempera+ure, Deq. Fahr 
Fig. 168. — Vapor pressures of various liquid fuels. 



wRT 

V 



per square inch pressure is given by the equation v 

15. 1 X 53.4 X 520 

14 7 x 144 = 197 cu. ft. approx. 

It is desirable that the fuel entering the inlet manifold should 
be entirely in the vapor form, either superheated or just saturated. 
This condition is necessary to obtain a homogeneous mixture and 
an equal distribution of the fuel to all the cylinders. The possi- 
bility of obtaining this condition may be determined by continu- 
ing the preceding calculation. If the air is saturated with 
heptane vapor at 60°F. its partial pressure, p a , will be 14.7 — 0.54 
= 14.16 lb. per square inch and the volume of the air at this 



FUELS AND EXPLOSIVE MIXTURES 



229 



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230 THE AIRPLANE ENGINE 

14.7 
reduced pressure will be ttt^ X 197 = 202 cu. ft. This quantity, 

and similar quantities for other air temperatures and other fuels, 
are given in Table 14. The vapor coexists in the same space with 
the air (202 cu. ft.) and as this volume is greater than the volume 
which the saturated vapor of heptane occupies (103 cu. ft.) the 
vapor cannot be saturated in a chemically correct mixture; the 
vapor will be superheated. At the temperature of 40°F. the air 
volume is seen to be 194 cu. ft. and that of the saturated vapor of 
heptane 185 cu. ft.; as these are approximately equal the vapor 
will be practically saturated. At temperatures lower than 40°F. 
the volume of the air will be less than the volume of the saturated 
vapor and in that case part of the fuel will necessarily be in the 
liquid form. An excess of air above that chemically necessary 
will lower the temperature at which liquid must begin to appear; 
an excess of fuel will raise that temperature. 

With a less volatile fuel such as octane it will be seen by 
inspection of the table that a higher temperature (a little under 
80°F.) will be necessary if the fuel is to be in the vapor form. 
With benzol the temperature is well below 40°; with methyl and 
ethyl alcohol between 70 and 80°F. 

It should be noted that the temperatures of the table are the 
temperatures after the vapor is formed. In the carburetor, the 
latent heat of vaporization of the fuel is taken from the air and 
the liquid fuel, with the result that the temperature of the mixture 
falls below the temperature of the entering air and fuel, unless 
heat, equal to the latent heat, is supplied from the jacket water or 
exhaust gases. If the latent heat of the fuel is 135 B.t.u., the 
specific heat of the liquid 0.45 and the specific heat of air at 
constant pressure 0.241, the fall in temperature AT for a mixture 
of 1 lb. of fuel with 15.1 lb. of air is given by the equation 

135 = A!T(0.45 + 15.1 X 0.241) 
or AT = 33°F. 

If the fuel is just saturated at 40°F., the entering temperature 
of the air and fuel would have to be at least 40 + 33 = 73°F. to 
permit all the fuel to be vaporized, if no heat is supplied to the 
mixture from outside. 

The following table 1 gives data of a similar nature for various 
fuels. Column 3 gives the temperature of the fuel-air mixture at 

1 From Kutzbach, Technical Note No, 62. National Advisory Committee for 
Aeronautics, 1921. 



FUELS AND EXPLOSIVE MIXTURES 



231 



which the vapor of the fuel is just saturated; the mixture is 
supposed to be chemically correct and the pressure of the mixture 
is atmospheric pressure. In column 4 is given the fall in tempera- 
ture of the air and liquid fuel required to supply the latent heat 
for complete vaporization. 1 The initial temperature of the 
air must be at least equal to the sum of the quantities given 
in columns 3 and 4; this sum is given in the last column. 
Saturation Temperatures of Air-fuel Mixtures 



Fuel 


Boiling 

point, 

deg. F. 


Saturation 

temperature 

of the fuel \ 

mixture, 

deg. F. 


Drop in 
temperature 

due to 
evaporation, 

deg. F. 


Minumum 
temperature 
of the air for 

complete 

vaporization, 

deg. F 


Hexane 


158 
176 
172" 
320£ 

428 




23 

72 

108 

198 


54 

54 

198 

63 

72 


54 


Benzene 


77 


Ethyl alcohol 

Decane 


270 
171 


Naphthalene 


270 



It is evident that the temperature of the air-fuel mixture with 
decane or naphthalene as fuel is so high as to reduce considerably 
the volumetric efficiency and the power of the engine if all the 
fuel enters in the vapor form. 

Gaseous Explosions. — In an airplane engine making 1,800 
revolutions per minute, the duration of the explosion should not 
be greater than the time of one-sixth of a revolution or J^80 
second. The possibility of employing a gasoline engine depends 
on the possibility of carrying out the explosion process with a 
high degree of completeness in this extremely short time. 

Explosion is a chemical reaction attended by the liberation of 
a considerable amount of heat. It is a combustion process. 
Combustion results from the chemical union of a fuel with oxygen 
and this union may take place either (1) at the place where the 
two are brought into contact as with the ordinary gas burner, or 
(2) in an intimate mixture of the two, as in a bunsen burner or 
in a gas engine cylinder. Explosive reaction can take place 
only with an intimate mixture. 

The reaction in an intimate mixture is not necessarily explosive; 
for example, no explosion occurs in the bunsen burner. An 

1 Some of these values are calculated by Kutzbach from values of latent 
heat which are apparently too high. 



232 



THE AIRPLANE ENGINE 



explosion is always self -propagating : that is, if part of the mixture 
is ignited the combustion will spread throughout the mass of 
the mixture. The term " explosion" is commonly reserved for 
the case where the velocity of such propagation is high; but there 
is no definite line of demarcation between explosion and slow 
burning. 

The velocity of propagation of combustion in an explosive 
mixture depends on the kind of fuel, the amount of oxygen 
present, the amount of inert gases present, the temperature, 
pressure, and a number of other factors. The strength of the 
explosive mixture is the most important factor. No explosion 
is possible if the ratio of air to fuel exceeds certain limits. Bunte 1 
has found the explosive limits for various air-fuel mixtures at 
atmospheric pressure and temperature as given in the following 
table : 

Explosive Limits op Air-fuel Mixtures 



Ratio of air to gas by 
volume 



Fuel 



Lower limit, 
air in 
excess 



Upper limit, 
gas in 
excess 



Theoretical 

ratio of air 

to gas by 

volume 



Carbon monoxide 

Hydrogen 

Water gas 

Acetylene 

Coal gas 

Ethylene.-. 

Alcohol 

Marsh gas 

Ether 

Benzene 

Pentane 



5.06 
9.58 
7.06 
28.8 
11.6 
23.4 
24.3 
15.4 
35.7 
36.7 
40.7 



0.33 
0.50 
0.49 
0.91 
4.23 
5.84 
6.32 
6.81 
12.0 
14.4 
19.4 



2.4 

2.4 

2.4 

11.98 

5.7 

14.4 

14.4 

9.63 

28.41 

36.0 

37.5 



Burrell and Gauger 2 give explosive limits of air-gasoline mixtures 
as 66 and 16 (ratio of air to gasoline vapor by volume). 

The above results were obtained with mixtures at ordinary 
atmospheric pressures and temperatures. They show that a 
self-propagating combustion is possible with most fuels where 

1 The Engineer, March 28, 1902. 

2 Technical Paper 150, U. S. Bureau of Mines. 



FUELS AND EXPLOSIVE MIXTURES 233 

there is a considerable excess present either of air or of fuel. 
These limits are considerably extended as temperature and pres- 
sure increase. For example, at 600°C. it is possible to explode 
a mixture of CO with 12 times its volume of air, as compared with 
5.06 times at atmospheric temperature. The presence of carbon 
dioxide in place of some of the excess air diminishes the explosive 
limits. 

The temperature to which part, or all, of the mixture must be 
brought to initiate an explosion is called the ignition temperature. 
This varies with the fuel, strength of mixture, the volume or mass 
of the mixture heated, the temperature and dimensions of the 
containing vessel, and the method of ignition. A weak spark, 
although it has a temperature much higher than the ignition 
temperature, may fail to cause an explosion. It may start 
combustion at the place where it passes, but the heat loss by 
convection, conduction and radiation may be in excess of the heat 
of combustion and the flame will fail to propagate. A sufficient 
duration of spark is also necessary. A flame may ignite a 
mixture that cannot be exploded by a spark, because it gives, 
initially, so large a volume of flame that the radiation loss to the 
containing vessel does not cool it below the ignition temperature. 
If the whole mass is raised in temperature simultaneously (as 
by adiabatic compression) the ignition temperature will be less 
than when part of the mixture only is heated. This ignition 
temperature, with adiabatic heating of fuel-air mixtures, is 
from about 1,200°F. for hydrogen to about 1,700°F. for carbon 
monoxide. With the usual gas engine fuels it falls between those 
limits, the value depending on the hydrogen and the neutrals 
present. 

The ignition temperature has great importance as it determines 
the permissible ratio of compression, and thereby, the limit of 
efficiency in the engine. Compression must stop just short of 
that temperature at which ignition will occur. Any means for 
increasing the cooling of the mixture during compression (such 
as improved water jacketing) will permit a greater ratio of com- 
pression. Local heating of the mixture, as by carbon deposit, 
may result in preignition. 

Combustion once started in an explosive mixture may either 
die out or be propagated. If it once starts to propagate itself, 
it is likely to continue and there will result an explosion. The 
velocity with which the combustion is propagated increases 



234 THE AIRPLANE ENGINE 

progressively in all true explosions. In the case of a bunsen 
burner the velocity remains constant and the combustion is not 
explosive. The flame in that case is stationary, but as the gas is 
moving the flame is really moving relative to the gas, in the 
opposite direction and with the same velocity. If the velocity 
of the gas is diminished too much by partly closing the gas 
supply, the flame will shoot back, i.e., the flame will travel more 
rapidly than the gas. The flame remains at the mouth of the 
burner under considerable variations of gas velocity in the 
burner because the velocity "of the mixture decreases rapidly as 
it issues from the burner, so that there will be some place, close to 
the burner, at which the gas velocity equals the velocity of flame 
propagation. The flame will remain stationary at that place. 
The cooling effect exerted by the metal burner also reduces the 
flame propagation velocity. 

If the velocity of the gas which will just keep the flame away 
from the burner is measured, it will give a rough indication 
of the velocity of flame propagation in the mixture. The results 
will not be very accurate because of cooling and diluting influ- 
ences of the atmosphere. Experiments of that general nature 
show that, at atmospheric temperature and pressure, for H and 0, 
the velocity of propagation is about 115 ft. per second, and for 
CO and O about 4J^ ft. per second. This is for the combining 
proportions, which give approximately maximum velocities. 
With H and air the velocity drops to about 10 ft. per second 
at 212°F. For gasoline-air mixtures, at atmospheric temper- 
atures, velocities of about 3.5 ft. per second and for alcohol about 
3 ft. per second are realized. These results apply only to linear 
propagation at atmospheric pressure. 

In a closed vessel, such as a gas engine cylinder, the conditions 
are quite different. The propagation, starting from a point, is 
spherical; the increase of temperature results in increase of 
pressure and as the flame spreads the unburned portion will 
be compressed adiabatically and will increase continually in 
pressure and in temperature. As the temperature increases the 
rate of propagation will increase. The velocity of propagation 
will then be continually accelerated. The flame, moreover, is 
carried forward bodily by the expansion of the burned portion. 

Experiments on explosions in closed vessels have determined 
the time required to reach maximum pressure with various 
mixtures exploded in vessels of various shapes. If the maximum 



FUELS AND EXPLOSIVE MIXTURES 235 

distance from the ignition point to the boundary of vessel is 
divided by this time, the quotient gives a measure of the average 
rate of flame propagation. With illuminating gas at atmospheric 
temperature, in a tube J^ m - m diameter and with 73^ in. travel 
of flame, this varies from 5 to 24 ft. per second, according to 
the strength of the mixture. It increases rapidly with increased 
initial temperature; in some cases as the tenth power of the 
absolute temperature (= 1,000-fold for doubled temperature). 

With the largest existing gas engines (using blast-furnace gas) 
the available time for a good explosion is about }■$ sec. and the 
maximum distance the flame must travel is about l}i ft.; this 
gives a mean velocity of 11 ft. per second. The addition of a 
third igniter has sometimes increased the capacity 20 per cent and 
shows that the speed limit has been reached. Blast-furnace gas 
consists mainly of CO, which, at low temperature, has a velocity 
of propagation not greater than one-third that of gasoline. 

With an airplane engine at 1,800 r.p.m., the time for explosion 
is about Hso second; if the flame travels 2 in. the mean velocity 
will be 33 ft. per second. By increasing the ignition lead, still 
more time might be provided; the speed of the airplane engine is 
not yet limited by the velocity of propagation of the explosion. 
Alcohol is slower so that alcohol engines could not be run as fast 
as gasoline engines if the rate of propagation of the explosion 
should ultimately determine the limit of speed, instead of valve 
areas and inertia effects as at present. 

The observed velocities of propagation in actual engines are 
higher than those which experiments with closed vessel indicate. 
This results from another factor, turbulence. The velocity with 
which the gases enter gas-engine cylinders is very much higher 
than the velocity of propagation of flame. With 1-lb. drop of 
pressure into the cy Under and no frictional resistance the velocity 
of the entering air would be about 350 ft. per second; with J^ lb., 
175 ft.; with }{ lb. about 120 ft. per second. This gas velocity 
causes turbulent conditions which cannot be quieted down by the 
time explosion starts. The propagation is not spherical but is 
by currents and eddies of burning gas which carry flame to all 
parts of the vessel more rapidly than is possible with spher- 
ical propagation. Recent experimental work bears this out. 
Dugald Clerk found, in a common gas-engine cylinder, that after 
quieting down turbulence, the explosion takes nearly three times 
as long as when the usual conditions exist. Experiments in 



236 THE AIRPLANE ENGINE 

closed vessel without stirring gave the time of explosion as 0.13 
sec; with vigorous stirring the time required was only one-sixth 
as long. 

Detonation. — During explosion in a closed vessel the advancing 
flame sphere sends off compression waves which travel through 
the unburned mixture with the velocity of sound in that medium. 
If the vessel is of sufficient dimensions the increasing velocity of 
the flame and the continuously increasing pressure and tempera- 
ture of the unburned mixture will result in the formation of a 
wave in which the pressure will be such as to bring the mixture 
(adiabatically) to the ignition temperature. In that case the 
wave will cause combustion as it moves on. The velocity of this 
wave will be greater than that of sound because the process is 
not merely one of wave transmission but of chemical reaction 
also. Investigations of the explosive wave show velocities of 
the order of magnitude of 3,000 to 6,000 ft. per second and 
pressures of 1,000 to 2,000 lb. per square inch. These pressures 
are destructive to engines and should be avoided. 

The " detonations" or " pinking" which are both felt and 
heard in engine cylinders under certain conditions of operation 
probably indicate either the generation of an explosive wave or 
breaking down of the fuel with the liberation of free hydrogen, 
which explodes with extreme rapidity. In such cases the com- 
bustion is notably incomplete, the exhaust containing much 
free carbon, and the power and efficiency of the engine fall off. 
Fuels consisting of paraffins have a low-ignition temperature, 
and are readily detonated. Fuels belonging to the aromatic 
group have higher ignition temperatures and can be used with 
higher compression pressures without detonation. 

The maximum pressures to which fuels can be compressed 
without serious detonation have been determined by Ricardo, 1 
who used for that purpose a variable compression engine with 
compact combustion space, central igniter, and other features 
making for maximum capacity and efficiency. His results, 
including the corresponding indicated mean effective pressures 
and thermal efficiencies, are given in Table 15. The data for 
toluene, xylene, and acetone are for a compression ratio of seven, 
which gives a compression pressure well below their detonation 
pressures; it was not considered desirable to go above that com- 
pression ratio for hydrocarbon fuels on account of the excessive 

1 The Automobile Engineer, Jan. and Feb., 1921. 



FUELS AND EXPLOSIVE MIXTURES 



237 



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238 



THE AIRPLANE ENGINE 



explosion pressures reached. Bicardo concludes from his investi- 
gations that detonation is less the lower the rate of burning of the 
fuel, and that no fuel has a rate of burning too low to permit of 
maximum efficiency being obtained at the highest practicable 
engine speeds. All the hydrocarbon fuels give the same power 
output and efficiency, within 2 per cent, for the same ratio of 
compression so long as that compression is not high enough to 
produce detonation. Figure 169 gives the indicated mean 
effective pressure and indicated thermal efficiency practically 
attainable in a high-grade engine with a non-detonating hydro- 
carbon fuel. With alcohol the high latent heat permits the use of 
a greater weight of charge and consequently of increased output 
as compared with the hydrocarbons. 1 




3.5 



4.0 4.5 5.0 5.5 6.0 6.5 
Ratio of Compression. 



7.0 



Fig. 169.- 



-Maximum attainable m.e.p. and thermal efficiency in Otto cycle 
engine using hydrocarbon fuel. 



The performance of a mixture of hydrocarbons is found to be 
the mean performance of the components. The highest per- 
missible compression pressure is determined by the relative pro- 
portions of aromatics, naphthenes and paraffins; the smaller the 
proportion of the last the higher may be the compression pressure. 
The heavier compounds in the paraffin series detonate at lower 
compression pressures than the lighter compounds. 

The tendency of a fuel to detonate is expressed by Bicardo in 
terms of its "toluene value." The scale of toluene values is 
based on compression pressures at detonation; it is per cent for 
a selected standardized gasoline detonating at a compression ratio 
of 4.85, is 100 per cent for toluene, and varies in direct propor- 
tion to the change in compression pressure. Its value is negative 
for fuels which detonate at lower compression pressures than 
the standard gasoline. 

1 The vaporization takes place mainly in the cylinder, during admission 
and compression, and keeps down the compression temperature. 



FUELS AND EXPLOSIVE MIXTURES 



239 



The influence on the detonating temperature, and consequently 
on the permissible compression pressure, of the addition of 
toluene to a paraffin fuel is shown by the following tests. 1 
Toluene, percent.... 0.0 10.0 20.0 30.0 40.0 50.0 60.0 
Compression ratio... 4.85 5.20 5.57 5.94 6.32 6.67 7.05 
Indicated m.e.p...... 132. 5 135.4 138.7 142.0 144.9 147.5 150.0 



420 



400 



u 360 

o 

v 360 



340 



320 



300 



280 



260 





Brake- M.ER 














— k 


::::: ^: 


"^^ 






~*«* 
























• 


/ 






















A 


/ 












// 

/A 












/ 

/ 

// 




Per 

Cat 


ma.68°t 
if. 59° t 


te Sasol 
'e 6asol 


ine 
ine 































1Z0 



no 



100 



1700 



1800 



1000 1200 1300 1400 1500 1600 
Revolutions perMin. 
Fig. 170. — Variation of power of 12-cylinder Liberty engine with fuel. 

The effect of the addition of ethyl alcohol is even more marked 
than that of toluene; only three-fifths as much alcohol need be 
added to obtain the same compression ratios. The effect of these 
additions on engine capacity are indicated by the mean effective 
pressure values of the table; the increase in engine efficiency can 
be seen from Fig. 169. 

The preponderating importance of the detonating character- 

1 Ricabdo: Proc. Royal Aeronautical Society, 1920. 



240 



THE AIRPLANE ENGINE 



istics of a fuel has received more recognition in England than in 
this country. Fuels are blended with toluene, benzol or other 
aromatics or naphthenes so as to give a standard toluene value. 

This amounts to giving a 
standard detonating com- 
pression pressure when the 
fuel is used in a standard 
engine. The actual deto- 
nating pressure and there- 
fore the permissible 
compression ratio is largely 
determined by the charac- 
teristics of the engine in 
which the fuel is used. 
With a poorly shaped com- 
bustion space, non-central 
ignition, and other un- 
favorable features, a fuel 
will detonate at a much 
lower compression ratio 
than when the conditions 
are favorable. In such an 
engine, detonation may be 
prevented by the use of 
overrich mixtures and late 
ignition, with a resulting 
sacrifice of both power and 
economy. 

Influence of Fuel on 
Capacity. — A comparison 
of the power output of a 
Liberty 12 with two grades 
of gasoline is shown in 
Fig. 170; the low-grade 
(59°Be.) gasoline falls off 
rapidly in brake mean effective pressure above 1,500 r.p.m., while 
the 68°Be. gasoline maintains its mean effective pressure well to 
1,800 r.p.m. 

A comparison of six fuels is shown in Fig. 171. These fuels 
were used in a single-cylinder Liberty engine. Their distillation 
curves are given in Fig. 172; these represent about the full range of 




1200 1400 1600 1800 

Revolutions per Mm. 

Fig. 171. — Variation of power of single-cyl- 
inder Liberty engine with fuel. 



FUELS AND EXPLOSIVE MIXTURES 



241 



commercial airplane fuels. It will be seen from Fig. 171 that the 
total range of power is 2.8 h.p. at 1,800 r.p.m. with a maximum 
value of 37 h.p. at that speed; this power range is only 7.6 per cent. 




First 10 

Dro P Percentage Distilled 

Fig. 172. — Distillation curves of the fuels of Fig. 171. 

The tests of the Bureau of Mines 1 show the comparatively 
small range in efficiencies resulting from the use of different 
fuels. The fuels include straight refinery, blended casing-head, 

Calorific Value, Power Developed in Engine Tests, and Specific 

Gravity of Various Typical Gasolines from Mid-Conttnent 

and Eastern Fields 



Field from which 

sample was 

obtained 


Process of 
manufacture 


Gravity 


High calorific 
value of gasoline 


Power 
developed, 
horse- 
power 
hours per 
pound of 
gasoline 


Specific 
gravity 


Be. 


Calories 

per 

gram 


B.t.u. 

per 
pound 


Mid-Continent 

Mid-Continent 

Mid-Continent 


Cracking plant 
"Straight" refinery 
"Straight" refinery 
"Straight" refinery 
"Straight" refinery 
"Straight" refinery 
Blended casing-head 
"Straight" refinery 
"Straight" refinery 
"Straight" refinery 


0.745 
0.742 
0.733 
0.718 
0.724 
0.727 
0.733 
0.724 
0.715 
0.687 


57.9 
58.7 
61.0 
65.0 
63.4 
62.6 
61.0 
63.4 
65.8 
73.8 


11,165 
11,174 
11,180 
11,187 
11,215 
11,221 
11,230 
11,236 
11,250 
11,315 


20,097 
20,113 
20,124 
20,137 
20,187 
20,198 
20,214 
20,225 
20,250 
20,367 


1.345 
1.403 
1.350 
1.405 


Mid-Continent 

Mid-Continent 


1.395 
1.396 
1.376 




1.420 


Mid-Continent 


1.365 

1.487 







and cracked gasoline with densities varying from 57.9 to 73.8°Be. 
The accompanying table shows that the work done per pound of 
1 Technical Paper 163, U. S. Bureau of Mines. 



16 



242 THE AIRPLANE ENGINE 

gasoline varied from 1.345 to 1.487 h.p.h., the higher value being 

generally obtained with the lighter gasolines. The heats of 

combustion also increased slightly as the gasoline became lighter, 

but the total range of higher heat values is only slightly greater 

than 1 per cent. Neglecting this, and assuming alow heat value 

of 18,500 B.t.u. for all the fuels, the range of efficiencies is seen 

. ■, 1.345 X 33,000 X 60 1Q , K .. 

to be from 18 500 V 77 8 = 18.45 per cent to 20.48 per 

cent. The engine on which the tests were made is of low com- 
pression and low efficiency; the indications are that the change in 
efficiency with the fuel is even less in engines of higher efficiency 
such as are used in airplane practice. 

A mixture of alcohol and gasoline has been used in high- 
compression aviation engines. This mixture eliminates deton- 
ation and has good starting characteristics. An "alcogas" 
tested at the Bureau of Standards 1 contained 40 per cent alcohol, 
35 per cent gasoline, 17 per cent benzol and 8 per cent of other 
ingredients. With a compression ratio of 5.6, the maximum 
power at ground level was the same as for a high-grade aviation 
gasoline, but as the level increased the power output became 4 to 
6 per cent greater than for gasoline. The thermal efficiency was 
superior by about 15 per cent; the fuel consumption was increased 
on account of the lower heat value per pound of fuel. At a 
compression ratio of 7.2 the power output was increased by 
about 16 per cent as compared with gasoline at 5.6 compression 
ratio and the thermal efficiency increased about 22 per cent, which 
just offsets the lower heat of combustion of the fuel and gives 
the same weight of fuel per brake horse-power hour for both fuels. 
The distillation curve of the alcogas is very flat, 80 per cent distill- 
ing off between 140 and 175°F.; the end point is high, 360°F. 

EXPERIMENTAL DETERMINATION OF STRENGTH OF MIXTURE 

The accurate determination of the ratio of air to fuel used by 
an engine requires the separate measurements of the weights of 
fuel and air. The weight of fuel is readily ascertained, but 
the weight of air offers difficulties. The simplest method is 
to connect the air intake of the carburetor, in an airtight manner, 
with a large box to which air is admitted through a standard or 
calibrated sharp-edged orifice. If there is more than one air 
intake it is better, if practicable, to enclose the whole carburetor 

1 National Advisory Committee for Aeronautics, Report No. 89. 



FUELS AND EXPLOSIVE MIXTURES 



243 



in an airtight chamber, connected with the orifice box. The air 
should enter the orifice quietly, passing through a honeycomb 
screen to eliminate the effect of wind or air currents. 

The weight flowing through such an orifice is given by the 
equation 

M = l.lFy[$(p -p.)XC 

where, M is the weight of air flowing per second in pounds. 
F is the area of the orifice in square inches. 
p is the external atmospheric pressure, pounds per square 

inch absolute. 
T is the atmospheric temperature, degrees absolute, 

Fahrenheit. 
p is the pressure inside the orifice box, pounds per square 

inch absolute. 
C is a constant 

The value of C has been determined by Durley 1 with great 
accuracy for. circular sharp-edged orifices in steel plates, 0.057 
in. thick. The following' table gives his values for the coeffi- 
cient C: 

Coefficients of Discharge for Sharp-edged Orifice 





Pressure difference on 


two sides of orifice, inches of water 


Diameter of 












orifice, inches 














1 


2 


3 


4 . 


5 


He 


0.603 


0.606 


0.610 


0.613 


0.616 


y 2 


. 0.602 


0.605 


0.608 


0.610 


0.613 


1.0 


0.607 


0.603 


0.605 


0.606 


0.607 


2.0 


• 0.600 


0.600 


0.600 


0.600 


0.600 


3.0 


0.599 


0.598 


0.597 


0.596 


0.596 


4.0 


0.598 


0.597 


0.595 


0.594 


0.593 


4.5 


0.598 


0.596 


0.594 


0.593 


0.592 



It will be observed that the coefficient is constant for a 2-in. 
orifice. 

In most cases it will not be found practicable to measure the 
air in this manner. A good approximation can be obtained from 
a measurement of the pressure drop from the mouth to the 

1 Trans. Am. Soc. Mech. Eng. f 1906. 



244 



THE AIRPLANE ENGINE 



throat of the choke or venturi tube of the carburetor. This 
drop can be obtained by connecting a water manometer with a 
very small hole (^2 in.) drilled into the smallest section of the 
choke. Tests carried out on a considerable number of carbu- 
retors show that the coefficient of discharge, C, varies only 
slightly with the form and dimensions of the venturi tube. The 
weight of air flowing through a carburetor of F sq. in. free area 
at the throat is given by the equation 

where M, p and T have the same meanings as for an orifice and 
p is the pressure at the throat. The coefficient C varies from 
0.82 to 0.85 and may be assumed to have the mean value 0.84. 

















— 1 — 1 — 1 
c o 2 






































































































s 


/ 
























s 














<p s 










^d 







s 































































n 



ZO IS 16 14 12 

Ratio of Air to Gasoline 



Fig. 173. — Composition of the exhaust gases from a gasoline engine. 

Still another method is available if actual air and fuel measure- 
ments are impracticable. The investigations of Watson, on 
automobile engines, have shown that the composition of the 
exhaust gases varies in a regular manner with the strength of 
the mixture of air and gasoline admitted to the cylinder. His 
results are shown graphically in Fig. 173. With a chemically 
perfect mixture of about 14.5 parts of air to one of gasoline the 
exhaust gases contain about 13 per cent of C0 2 by volume and 
about 0.5 per cent each of O2 and CO. If the air is present in 
excess (weaker mixture) there is more free 2 and less C0 2 in the 
exhaust; if the mixture is richer, the free 2 disappears and the 
amount of CO increases while the C0 2 decreases. All that is 
necessary for the test is an Orsat or other volumetric gas-analysis 
apparatus and the determination of the C0 2 and 2 content, or 
in case no 2 is present, the C0 2 and CO content. 



CHAPTER X 



THE CARBURETOR 



7c? Engine 



An ideal explosive mixture arriving at the intake manifold 
of an engine should have the following characteristics: (1) 
it should be homogeneous throughout, (2) it should be of the 
composition or strength to develop maximum economy under 
each condition of engine operation, and (3) it should permit of the 
development of the maximum possible power. 

In a stationary constant-speed engine, in which engine torque 
alone is variable, these results might be approximated by the 
use of an injection valve, under the control of the governor, 
spraying finely atomized fuel into the current of air going to 
the cylinders. In an automobile engine, with both engine torque 
and speed variable, this simple injection method cannot give 
satisfactory results. In the airplane engine with the three 
main variables of torque, 
speed and air density, the 
problem is even more com- 
plicated. For such engines 
the explosive mixture is 
formed by the use of a 
carburetor. 

A carburetor is a device 
in which part or all of the 
air going to the engine 
passes through a restricted 
passage, thereby acquiring 
velocity with consequent 
fall of pressure; the fuel is 
sucked into the current of 
air in an amount which 
varies with the pressure 

drop. In the simplified standard form of carburetor shown in 
Fig. 174, air flows through the restricted " choke," C, and creates 
a partial vacuum. Gasoline is maintained at a constant level in 
the float chamber by the action of the float, F, which controls the 
position of the needle valve, V, past which the gasoline enters. 
As the float chamber is open to the atmosphere the level of 

245 




V 

Gasoline 
Fig. 174. — Diagram of simple carburetor. 



246 THE AIRPLANE ENGINE 

gasoline in the nozzle or jet, J, will be the same as that in the 
float chamber so long as the engine is not operating. The dis- 
charge orifice of the nozzle is placed higher than the gasoline 
level in the float chamber to prevent overflow of the gasoline 
into the air passage when the engine is standing in such position 
as to incline the carburetor at a moderate angle to the position 
shown in the figure. When air is drawn through the carburetor, 
increasing reduction of pressure at C, resulting from increasing 
velocity of the air, will give an increasing head on the gasoline 
and will cause an increasing weight flow of the fuel. The 
mixture of air and fuel will be of constant strength if the weight 
of gasoline discharged by the jet is directly proportional to the 
weight of air flowing through the choke. The actual strength 
of the mixture whether constant or not is controlled by the size 
of the gasoline jet. 

A carburetor built as in Fig. 174 would not discharge a mix- 
ture of constant strength for all rates of air flow, nor is such 
constancy desirable. It is common experience that the mixture 
delivered to the engine should be richer at very light loads 
(idling) than for heavier loads, and also that it should be richer 
for maximum power than for maximum economy. A satis- 
factory carburetor should vary the strength of the mixture so as 
to maintain the desired strength under all conditions of operation 
of the engine. 

A study of the action of a carburetor requires a knowledge of 
the laws of flow of gases and liquids through such passages as 
are found in carburetors. The more important results of 
experiment on such flow are given in the following pages. 

Theoretical Flow of Air through a Constricted Tube. — When 
air is flowing steadily through a tube whose cross-section varies, 
the weight and the total energy passing each section of the tube 
per second are constant. 

Let W = Weight of air passing in pounds per second. 

p = The air pressure in pounds per square foot absolute. 
P = The air pressure in pounds per square inch absolute. 
v = The specific volume of the air in cubic feet per pound. 
V = The velocity of the air in feet per second. 
T = The absolute temperature of the air in degrees Fahrenheit. 
/ = The internal energy of the air per pound in foot-pounds. 
A = The cross-section of the tube in square feet. 
a = The cross-section of the tube in square inches. 
q - Gravitational acceleration = 32.16 feet per second per second. 



THE CARBURETOR 



247 



The total energy passing any cross-section with unit mass of air 
is the sum of the internal energy I, the displacement work pv, 
and the kinetic energy V 2 /2g at that section. Assuming no 
heat transfer through the tube the total energy at 1 (Fig. 175) 
can be written equal to that at 2. 



TV y 2 2 

Ii + Vi vi + 2~ = h + P2V2 + -^r 



Fig. 175. — Venturi tube of optimum proportions. 



(i) 



Air is practically a perfect gas. If the expansion is without 
eddies or friction and without transfer of heat (adiabatic) 

P1V1 — P2V2 ' . 



h-h = 



n - 1 



(2) 



_ specific heat at constant pressure _ 
— specific heat at constant volume 
Furthermore, with adiabatic expansion 

PlVi n = P2V2 n (3) 

Substituting from equations (2) and (3) in equation (1) there 
may be obtained the equation 



TV 



TV 



-iS^t 1 -©"] 



(4) 



2flf 2<7 

The weight flow past any section is equal to the volume passing 
that section per second divided by the specific volume, or, 

V 

Since the weight flow is constant at all sections 

V1A1 7 2 A 2 



W 



Vi 



V2 



(5) 



and substituting from equation (3) 

Vi Vi 



•©" 



7, = V 



A 2 /p 2 \" 

'■aApJ 



(6) 



248 



THE AIRPLANE ENGINE 



Substituting this value of Vi in equations (4) and (5) 



y 2 = 



and 



W 




\pi/ \ y n - 1 vi 



(£)'& 



1 



If the section A\ is taken just outside the tube where the cross- 
section may be regarded as infinite and the air velocity zero, 
these last equations become 



and 






(9) 



=^^€©"-©"1 (10) 

In the use of equation (10) the specific volume, v h has to be deter- 
mined from the known pressure, p h and absolute temperature, T u 
by the gas equation, pm = RTi = 53.347Y Furthermore, 
it is practically most convenient to deal with pressures in pounds 
per square inch, P, and with areas in square inches, a. Substitut- 
ing the numerical values of g and n, substituting P and a for p 
and A, and substituting 53.34Ti/pi for v h equation (10) becomes 

This equation is difficult to use when the desired weight flow, W, is 
known and the pressure drop is required. The curves of Fig. 176 
are based on this equation. The ordinates are weight flows per 
square inch of area per minute, and the abscissae are pressure 
drops measured in inches of water. One pound per square inch 
equals 27.70 inches of water. An initial temperature of 60°F. 



THE CARBURETOR 



249 



(520° absolute) is assumed. The separate curves are for the 
initial pressures, Pi, marked on them. 

Equations (8) and (10) have a limit to their range of applica- 
tion. If air, initially of pressure pi and specific volume Vi, flows 
through a tube the smallest section of which is A 2 , the weight 
flow varies with the pressure, p 2 , at that section. The weight 

20 



18 



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20 40 60 80 

Pressure Drop in Inches of Water 

Fig. 176.— Weight flow of air. 



100 



120 



flow will be found from the equation to reach a maximum value as 
p 2 diminishes to a certain critical value, and then will apparently 
diminish as p 2 is still further reduced. This critical pressure 
occurs when 

- = (-^^=0.53 (for air) 



250 THE AIRPLANE ENGINE 

That is, the maximum weight flow will occur when the pressure at 
the smallest cross-section of the tube is 53 per cent of the initial 
or maximum pressure. It is found by experiment that the 
pressure at the smallest cross-section is never less than this 
amount, and that it remains at that exact value so long as the 
pressure on the downstream side is equal to or less than that 
pressure. The weight flow through a frictionless tube is deter- 
mined by the area of the smallest cross-section and cannot be 
increased by decreasing the pressure on the downstream side of 
that section below the critical pressure. In equations (8) and 
(10) t?2 can never have a value lower than 0.53pi. 

For very small pressure ranges, for example when (pi — p 2 ) is 
equal to or less than 1 per cent of p h the expansion of the air 
resulting from the pressure drop is so small as to be negligible 
and the flow may be assumed to follow the simpler laws of flow 
of incompressible fluids. In this case 7i = 7 2 , and v\ = v 2 , and 
equation (1) becomes 

V 2 2 7i 2 

-2g ~ ~2g = *** ~ ^ 

If the initial velocity is zero 

^ = (Pi " P2>1 (12) 

or V is proportional to \/pi — p 2 , and since the weight flow is 
proportional to V (with constant cross-section and constant air 
density), 

w = Ky/pT=pi (13) 

With air at atmospheric pressure, the error resulting from the 
use of this equation would be about —2.3 per cent for 1 lb. per 
square inch pressure drop, and is roughly proportional to the 
pressure drop for small pressure drops. Equation (12) shows 
that the velocity and therefore the weight of air flowing is pro- 
portional to the square root of the pressure drop so long as the pres- 
sure drop is small. 

The actual flow of air through a constricted tube is found to 
be less than the amount indicated by equation (10). Actual 
flow is always accompanied by frictional resistance and the 
formation of eddies. The ratio of the actual flow to the theo- 
retical flow of equation (10) is called the coefficient of discharge 
of the tube and its value has to be determined by experiment. 

The choke of a carburetor is usually of the general form shown 



THE CARBURETOR 251 

in Fig. 175, that is, it consists of three parts: (1) a converging 
entrance; (2) a throat; and (3) a diverging discharge; such a 
tube is generally described as a venturi tube. With equal areas 
at the entrance and exit, the pressure drop from the entrance to 
the throat would be entirely regained at the exit, if the air flow 
were frictionless and eddyless. In actual carburetors, the 
pressures at discharge will be less than that at entrance, and the 
difference will depend on the velocity of the air, the " stream- 
lining" of the passage, the degree of obstruction offered by the 
gasoline jet, and the weight of gasoline carried by the air. The 
total pressure drop in the venturi is important in determining 
the volumetric efficiency and capacity of the engine; to develop 
maximum power the charge should enter the cylinder with the 
maximum possible density. Loss of pressure in the carburetor 
is a direct source of loss of power in the engine. 

The results of published tests on comparatively large venturi 
tubes, with straight axes, and without obstruction at throat or 
entrance, show discharge coefficients varying from 0.94 to 0.99 
for cases where A\ = Az, and A 2 is equal to or less than 0.5 A\, 
In these tubes it is found that, for minimum friction and eddy loss, 
the included angle for the converging entrance should not exceed 
30 deg., and the diverging discharge tube should have an included 
angle between 5 deg. and 7.5 deg.; these should be joined to a 
short cylindrical throat by well rounded junctions. Figure 175 
shows a venturi tube of these optimum proportions. 

Such optimum proportions are generally not practicable for 
airplanes. Considerations of space available make it necessary 
to modify the entrance by curving its axis, and force the adoption 
of larger included angles. Furthermore, the air passage is 
obstructed by the gasoline jet and its supporting bosses, and, in 
many cases, by the throttle valve. All these factors will cause 
a diminution in the discharge coefficient and an increase in the 
pressure loss. An investigation at the Bureau of Standards 1 
gives data on certain carburetors which were designed for the 
Liberty engine. The air passages of these carburetors are 
shown in Fig. 177. The tests were made with various air 
densities (corresponding to different altitudes), and both with and 
without fuel admission. Figure 178 shows the coefficient of 
discharge; Fig. 179 the ratio of the exit to the entrance pressure, 

1 P. S. Tice: National Advisory Committee for Aeronautics, 4th annual 
report, pp. 608-615. 



252 



THE AIRPLANE ENGINE 



for both carburetors, with air of 750 mm. pressure and with 
various weights of air flowing. Figure 180 shows the pressure 
recovery ratio for the Zenith carburetor, with various air densities, 
and both with and without fuel admission to the air. The 
conclusions derived from these tests are as follows : 





B 

Fig. 177. — Zenith (A) and Stewart- Warner (B) carburetors. 



1. The coefficient of discharge for the carburetor passages 
tested has an almost constant and maximum value for effective 
throat velocities greater than about 150 ft. per second. 



0.9 



0.6 











Carburetor A 










r. 
























Carb 


jretor 


■B 








\ 






















\ 



































































20 30 40 

Pi -P 2 (Inches of Water) 



50 



Fig. 178. — Venturi discharge coefficients for Zenith (A) and Stewart- Warner 

(B) carburetors. 

2. The value of the coefficient of discharge for the carburetor 
passages tested lies between 0.82 and 0.85, under service con- 
ditions. These values are probably typical of reasonably well 
formed passages of similar type. 



THE CARBURETOR 



253 



3. The coefficient of discharge for carburetor passages of this 
type is apparently only slightly modified as a result of consider- 



100 



;o.99 



£0.9 8 



mi 



0.96, 























































N& 






















N?S 


<5> 






































































\ 



























0.0c 



0.04 0.08 0.12 0.16 

Air- Lb. per Sec. per Sq. In. of Throaf Area 



0.20 



Fig. 179. — Pressure drops at partial loads in Zenith (A) and Stewart- Warner 

(B) carburetors. 

able changes in passage form, with respect to angles of entrance 
and exit. 

1.00 



0.99 



'0.91 



0.96 















A = Air onlu 
B=Air >> 


af 750mm 

» 550 » 








\ 


\ N. 


\ 




C=A;r » 

D= Air and fuel at 750 " 

E=Air » » » 550 » 








\\ 

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% N. 






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N 


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V 




\ 
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\ 












\ 






















\ 















0.02 



0.04 0.08 0.12 0.16 0.20 

Air - Lb. per Sec. per Sq. In. of Throar Area 



Fig. 180. — Pressure drop through a Zenith carburetor as affected by air density 
and the injection of fuel. 



4. The coefficient of discharge for a carburetor passage is 
practically unaffected by wide variations in atmospheric density 



254 THE AIRPLANE ENGINE 

(less than 1 per cent maximum variation between the density 
limits of 0.075 and 0.035 lb. per cubic foot). 

5. The coefficient of discharge for a carburetor passage is 
practically unaffected by the introduction of fuel to the air 
stream (fuel discharge introduces irregularities not to exceed 
plus or minus 1 per cent). 

6. The pressure loss in the carburetor outlet changes with 
the turbulence or internal motion of the air stream. 

7. The pressure loss in the carburetor outlet changes with the 
quantity of fuel admitted to the air stream, and with the method 
of dividing the fuel by spraying. 

Pulsating Flow. — The previous discussion relates to steady 
flow of air through the choke. In actual operation the flow is 
pulsating; each carburetor usually supplies three or four cylinders. 
With a maximum of four cylinders the carburetor will be supply- 
ing one cylinder only at any instant. The flow of the air through 
the carburetor is determined by the velocity of the piston in the 
cylinder to which the air is going. As this velocity is zero at the 
ends of the stroke and a maximum at midstroke, the variation in 
velocity of flow through the carburetor would be considerable 
were it not for the steadying effect of the intake manifold. 
The volume interposed between the carburetor and cylinder 
acts as an equalizing device and cuts down the pressure pulsations 
at the exit of the carburetor. Tests made in England and 
at the Bureau of Standards 1 show that for a given weight of 
air flowing under pulsating discharge the coefficient of discharge 
of the carburetor (as determined from pressure measurements 
at the throat), the pressure recovery ratio, and the strength of 
the mixture are practically the same as for steady flow. 

The Flow of Fuel through a Nozzle or Jet. — The flow of a 

liquid through an orifice is given by the expression V = C\/2gh , 

where V is the velocity of flow, C a coefficient, and h the head 

under which the flow is occurring. This expression becomes 

W = 60.2 Ca\/^h (14) 

where W = Weight of liquid discharged in pounds per minute. 

a = Area of passage in square inches. 

s = Specific gravity of the liquid (referred to water at 

60°F.) 
h = Head or pressure drop across the jet expressed in 
inches of water. 

1 National Advisory Committee for Aeronautics, 4th annual report, p. 616. 



THE CARBURETOR 



255 



The coefficient C includes losses due to skin friction, fluid friction, 
contraction, and end effects. Its value varies with the head, h, 
with change in shape of the entrance to the jet, with change in 
ratio of length, L, to diameter, D, of the passage, and with the 
viscosity of the fuel. 

Investigations by Tice 1 on the flow through jets show the in- 
fluence of these different factors on the value of C. The effect 
of the alteration of the shape of the jet entrance from square to 
chamfered is shown in Fig. 181. The diameter and- length 
are the same for both jets. The major effect of the chamfering 
is to reduce the contraction of the stream in the entrance, in 
this case, at heads above 2 in. in water. While the coefficient, 




EFFECT OF CHAMFERING ENDS 

OF PASSAGE 
Included ana le of chamfer = 60° 
Depth » » = 0.0081 

Diameter of passage = 0.040 
Length » » = 0.4016 J 

Length -± depth = 10.04 



12 16 ZO 

Head = h (Inches of Water) 



24 



28 



Fig. 181. — Discharge coefficients of square and chamfered jets. 

C, has considerably higher values with increase of h with the 
entrance chamfered in this way, it will be noted also that its 
value varies through wider limits. Chamfering has the very 
practical advantage in carburetor manufacture, that the angle 
and depth of the chamfer, within comparatively wide limits, 
have an almost negligible effect on the discharge; while, on the 
other hand, small departures from truth in the making of sharp 
square edges result in wide variations in the discharge. This, 
together with the great difficulty of producing duplicate parts hav- 
ing square edges free from burr, practically rules out the square 
edge for carburetor metering passages. 

Within the range of metering passage diameters used in general 
carburetor practice, it is found 'that the value of C increases with 
increase of D (Fig. 182). 

The effect upon C of change in the ratio L:D is brought out 
in Figs. 183 and 184. In the former, C is plotted against h for 

^Loc. cit., p. 603 



256 



THE AIRPLANE ENGINE 



several values of L:D with D a constant. In Fig. 184, C is 
plotted against L:D, each curve being representative of a 
constant value for h. 



Fig. 



1.0 



^0.8 



-»-0.6 



0.4 



0.2 













F 




























£ 














D 










, 


/ 










C 
A 
























1 1 1 
EFFECT 0N-C OF CHANGE IN-D 

L-^D SUBSTANTIALLY CONSTANT 

Submerqed Orifice-Chamfered Ends 
D" J L" L^D T°C J 




























A =0.0327 0.4041 12.36 23.25 
B* 0.0350 » 11.54 23 SO 


















C= 0.0373 n 10.84 24.10 1 
D= 0.0395 » 10.23 24.70 


















E = 
F = 


V.UJIU U.VI3V 

0.0357 

1 1 


0.420 24.45 

1 1 





12 16 20 24 

Head = h( Inches of Water) 



28 



Fig. 182. — Influence of diameter on the discharge coefficients of jets. 
1.0 



o0.8 

.c 
<n 

5 

^0.6 



,0.4 



0.2 















D^ 
























I 
















C ^ 
























A 


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A^ 














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y 


/*^ 


























/ 


s 










b 


FFECTOr or h m in i - 

WITH D CONSTANT 
jbmerged Orifice- Chamfered End 

D" L" L*D T°C 
= 0.0357 0.406 11.37 21.10 
= » 0.200 5.60 26.60 
= » 0.100 2.80 23.80 






/ 






























£ 

c 


















L 


= 


| 


0.015 

1 


• 0a 


12 2 

1 


4.40 







12 16 20 

Head = h (inches of Water) 



24 



28 



183. — Influence of the ratio of length to diameter on the discharge coeffici- 
ents of jets. 



P|,0 

D 



Q 

<^0i 

o 



06 



0.4 















~n 


1 1 

FFFFfT UPON 


1 1 1 

"f'OF CHANGF IM . 




1 


$>: 










VALUE OF "L-D'AT VARIOUS HEADS 
Submeraed Passaae -Chamfered 
















D = 0.0357" j | 

_l h=Head inW 


crier. 
















S9v 


























^ 






























-2 






























■-0.5 









14 



L/D = Length -J- Diameter 
Fig. 184. — Influence of the liquid head on the discharge coefficients of jets. 

A change in temperature, T, affects the discharge from a passage 
in two ways — through its influence on the density, s, and through 



THE CARBURETOR 



257 



the change in fluidity. For ordinary variations in T, the change 
in s is comparatively small and has very slight influence on the 
discharge. The curves A, B and C, in Fig. 185, for gasoline 
discharged from a jet at three temperatures, expresses the order 



„ i.o 

o 



0.6 



0.4 

















C V 


,B 














— 


—~- 




Jf 








i^^^-A 


£ Y 


,D 








y 


^ 


& 


^ 


►-=— 




EFFECT ON-C OF CHANGE IN : t' 
Gasoline Flowinq 


=— * 




Y 


^ 










Su 

A 


bmerqed Orifice -Chamfered End 

T>" L" L-D T°C 
= 0.0344 0.401 11.83 9.85 0. 

* » » " mo 0. 

= v » 23.80 0. 

Free Orifice -Square Ends 
= 0.042 0.005 0.119 24.50 
= 0.020 0.0.05 0.250 24.50 8 


5 

P 
755 




/ 


Y 












B 
C 


7 49 

139 


















D 
£ 


4.00 





J 12 14 16 

Head = h (Inches of Water) 



20 



24 



Fig. 185. — Influence of temperature on the discharge coefficients of jets. 

of magnitude of the effect upon C of change in fluidity resulting 
from change in T. These results are for a comparatively long 
passage, in which this effect is much greater than with the smaller 
values for L:D found in carburetor practice. The curves D and 




Fig. 186. 



60 80 100 

Temperature , Deg.C. 

-Variation of the fluidity of liquid fuels with temperature. 



E are for sharp-edged orifices; they show great constancy of C 
with variation both of h and of T: a change in T from 24.5°C. 
to 4°C. shows no appreciable change in C at any value of h. 



17 



258 



THE AIRPLANE ENGINE 



The fluidity of a liquid is the reciprocal of its viscosity. The 
variation of the fluidity of aviation engine fuels with temperature 
has been investigated by Herschel/ who finds the results shown 
in Fig. 186. The value of C for a jet will increase as the tem- 
perature and, therefore, the fluidity of the fuel increases. There 
is no fixed relation between the densities and fluidities of different 
fuels; a change of fuel will ordinarily result in a change in C. 

Mixture Characteristics of a Carburetor with Constant Air 
Density. — It has been shown by equation (13) that, for moderate 
pressure drops in the choke, the theoretical air flow, W, is sensibly 
proportional to the square root of the pressure drop, and with a 
constant coefficient of discharge this means that the actual 
air flow follows the same law. It has been further shown that 
with discharge through a sharp-edged orifice the flow of the 
fuel follows the same law. Consequently, it would seem possible 
to construct a carburetor in which the air-fuel ratio would remain 
constant for moderate air flows. Actual carburetor constructions 
do not, however, employ sharp-edged orifices on account of the 
production difficulties already mentioned. Furthermore, the air 
flow does not increase as rapidly as the square root of the pressure 

drop — for example, in Fig. 
176, with air initially at 
14.7 lb. per sq. in. pressure, 
as the pressure drop in- 
creases from 10 in. to 40 
in. of water, the weight 
flow instead of doubling 
increases only from 6.56 to 
12.6 or 1.92 times. At the 
same time, using the stand- 
ard form of chamfered jet, 
as shown in Fig. 181, the 
coefficient of discharge in- 
creases and thereby in- 
creases the flow of fluid more than two fold. The mixture will 
therefore increase in richness as the load increases. To offset 
this increase, the structure of Fig. 174 is modified in all com- 
mercial carburetors. These modifications are extremely diverse 
in character and can be such as to produce a constant mixture 



28 

26 

24 
o 

^22 
o 
a: 
_20 



10 



























, 






















\ 






















\ 
























k 
























j 












^ 


t£t£ '< 


'arni 
















^•'V 






>r 














--< 


f^ 


v& 


































1.0 



0.6 0.4 

Loads 



0.2 



Fig. 187. — Variation of mixture strength 
with load in Zenith, Stewart- Warner and 
Stromberg carburetors. 



1 Bureau of Standards Technologic Paper No. 125. 



THE CARBURETOR 259 

or almost any desired variation of mixture with load. Some of 
these constructions will be considered later. 

The results of tests on three special airplane carburetors at 
standard air density, shown in Fig. 187, are characteristic of the 
methods of variation of the air-fuel ratio with the load in actual 
carburetors. The Zenith carburetor shows a very constant mix- 
ture; the other two show enrichment of the mixture with dimin- 
ishing load, a characteristic exactly opposite to that of the simple 
carburetor of Fig. 174. The actual value of the air-fuel ratio 
depends on the size of the fuel orifice and is not characteristic 
of the type of construction. 

Mixture Characteristics of Carburetor with Variable Air 
Density. — When the air density changes, as a result of change 
of air pressure and temperature during the ascent of an airplane, 
a new disturbing element is introduced into the behavior of the 
carburetor. With level flight and wide-open throttle, the engine 
speed may be assumed, as a first approximation, to be constant at 
all altitudes; this is not the case since the engine speed may 
fall off as much as 10 or 12 per cent. The volume of air passing 
through the carburetor is equal (approximately) to the piston 
displacement of the engine per unit of time and may also be 
assumed to be constant. The weight of air, W, taken in will then 
be proportional to the air density, D. At any altitude x, 

W D 

where o indicates ground condition. With a sharp-edged fuel 
nozzle of constant coefficient of discharge, the weight of fuel dis- 
charged, w, is proportional to the square root of the pressure 
drop at the carburetor throat (equation 14) and this is, approxi- 
mately, proportional to the air density, D (equation 12). This 
may be written: 

Wo \D, 



o 



W 

If R is the air-fuel ratio — , then 

w' 



that is, the strength of the mixture varies inversely as the square 
root of the air density. As the air density is proportional to its 



260 



THE AIRPLANE ENGINE 



pressure and inversely as the absolute temperature, T, 
becomes 

Re 

Ra 



this 



'O _ Ll I X 

'X yl-Lx I O 



On going from the ground to an altitude of 30,000 ft., where the 
air density is 40 per cent of ground density, the air-fuel ratio 
would fall from 20 to 14, or the strength of the mixture would be 

— ^-p — -J X 100 = 43 per cent. 

The strength of mixture desired can only be determined by 
engine tests. Such tests have been carried out on Hispano- 
Suiza and Liberty engines at the Bureau of Standards. 1 The best 



cH40 
to 

S.120 

£ 1C0 
| 80 
% 60 

SB 40 

I 20 
s: 

-3 0, 













J>* 0.075 



















0=006 


5" 














<*— n- 




O-0.05 










td= 


W-ft 






o 


rt^o— 






















0=003* 




~° 


" 


° 


" 


D= 0.025 














„ 








M 


° 
































2 


! 


4 

Mi 


1 

xtut 




-e R 


1 
atic 


8 

) = A 


2 



-uel 


2 


2 



jj.140 
1120 
| 100 
| 80 

"I 60 
& 40 
| Z0 



k 




















>k 


^ k 






























^ 




























































P 
































Fu 


el C 


?nsu 


mpt!°2- 

r ' T 


"7f 









0.07 0.07 0.05 0.04 
Air Density in Lb. per Ca Ft. 



0.03 



2.4 c 
o 

2.21. 
E 

2.0 2 



I.6S 



1-4 £ 

1.2^ 
o 

I.OcS 



Fig. 188. — Influence of air-fuel ratio on 
brake m.e.p. at various air densities. 



Fig. 189. — Influence of air density on 
maximum power and maximum thermal 
efficiency. 



mixture to use depends on whether maximum power or maximum 
economy is wanted. The curves of Fig. 188 show how the brake 
m.e.p. varies with the mixture ratio at air densities, D, from 
0.075 to 0.025. As these results are for constant engine speed, 
they also show the method of variation of the horse power 
developed. It will be seen that maximum power, P, is obtained 
with an air-fuel ratio of 15 at all air densities. Maximum 
economy (minimum fuel consumption per brake horse-power 
hour) is obtained at the points crossed by the curve M; it is 
seen that at ground level (D = 0.075) the most economical 
air-fuel ratio is 23, and that this value diminishes (richness in- 
creases) as the air density decreases. The maximum economy 
curve runs very near to the limit of explosibility, this limit 
requiring an increasingly rich mixture as the compression pressure 

1 P. S. Tice: Nat. Adv. Comm. Aeronautics, Uh Annual Report, p. 624. 



THE CARBURETOR 261 

diminishes. The air-fuel ratio for maximum economy is given 
approximately by 

R = 106D + 15 

where D is the air density. In Fig. 189 the same data are re- 
plotted to show the variation of brake mean effective pressure, and 
of fuel consumption per brake horse-power hour, with the air 
density, both at maximum power, P, and maximum economy, M. 

It is not possible for a carburetor operating with wide-open 
throttle to give both maximum power and maximum economy 
without some kind of manual control since the demands for these 
two conditions differ only in the amount of fuel supplied. The 
condition of maximum economy alone is important, except for 
war purposes. For a flight of several hours' duration the com- 
bined weight of an engine and its fuel consumption will be less for 
a larger engine operating at maximum economy than for a 
smaller engine operating at maximum power and developing the 
same total power. The carburetor should be devised to give 
maximum economy at full throttle, with a manual control to 
increase the fuel supply so as to give maximum power if desired. 

Economy of operation at low loads is unimportant in heavier 
than air machines since this condition of operation is not possible 
for other than very short periods. In lighter than air machines 
the economy at low loads may be of more importance. At full 
throttle the most economical air-fuel ratio varies from 23 at the 
ground to 19 at half -ground density; for operation at partial 
loads these figures must be reduced. It is not desirable to 
operate an engine with the mixture giving maximum economy 
because this mixture is «o close to the limit of explosibility that 
slight changes in condition might result in exceeding that limit. 
Since the economy changes but slowly with change of mixture 
in the neighborhood of the optimum value, it is the practice 
to operate with smaller mixture ratios ; a value of 20 at the ground 
is seldom exceeded. 

The optimum mixture at partial loads may be presumed, 
as at full load, to be fairly near to the upper explosive limit. 
This limit changes with the load as a result of change in compres- 
sion pressure and temperature and of change in the percentage 
of diluting residual gases present. The compression pressure 
exerts considerable influence on the explosive properties of a 
weak mixture, necessitating the use of a stronger mixture as the 



262 



THE AIRPLANE ENGINE 



load diminishes. The temperature at the end of compression does 
not change much since the ratio of temperatures at the beginning 
and end of compression is a function of the ratio of compression 
which remains constant; it may be presumed that the tempera- 
ture effect is negligible. The effect of charge dilution on ex- 
plosibility has been investigated for mixtures of air with methane 
and with natural gas, the diluting agent being CO2. 1 Some of the 
results of this investigation are plotted in Fig. 190. It is seen 
that with 20 per cent CO2, a mixture of natural gas and air cannot 
be made to explode at atmospheric pressure and temperature; 



3" 

c 































^ 


r< 
























fe 


te. 
























fe 




































limit 


of e * 


olosik 


iijjy. 












L 


ow^r 

















Fig. 190. 



2 4 6 8 10 12 14 16 18 20 22 
Carbon Dioxide = Per Cervr by Volume 

-Influence of carbon dioxide dilution on the explosibility of a mixture 
of natural gas and air. 



as the percentage of C0 2 diminishes the upper and lower limits 
recede until with no C0 2 present we have the lower limit with 
5.2 per cent and the upper with 11.6 per cent of natural gas 
present. The figures for a gasoline-air mixture are probably not 
very different. With higher pressures and temperature the 
explosibility limits will be changed but the method of variation 
will be the same. 

The amount of dilution of the charge by residual gases can 
be calculated approximately if the temperature of these gases is 
assumed. The amount of such dilution will vary with the 
load since the residual gases fill the clearance at exhaust pressure 
and are approximately constant in weight at all loads. The 
amount of such dilution, d = W r /W c (where W r = weight of 
residual gases and W c = weight of fresh charge), is shown in 

Element: Bureau of Mines Technical Paper No. 43; "The Influence of 
Inert Gases on Inflammable Gaseous Mixtures." 



THE CARBURETOR 



263 



Fig. 191, l which also shows the corresponding compression 
pressures. These curves are for a ratio of compression of 5.5 
and must be regarded as approximations only. The pressure 
at the end of compression is well above atmospheric pressure 
and as the temperature is probably about 1,100°F. absolute the 
dilution can be carried further than indicated in Fig. 190 without 
exceeding the explosive limit. The pressure and dilution of the 
charge at partial loads are such as to demand a richer mixture 



140 



120 



100 



3 80 



60 



d-40 



20 

































Temp 


>eraf 


ure 




0SS& 


^ 






=3 


==5 


388 












\ 


















1 




\ 






































m\ 






\ 


V 










t 


$ i 






"> 














6j 






X 


N 


X 


2x 






J 


w 






**>i 


«:« 


&s- 


Vs - 


"^ 






// 








--- 




^~ 






















// 


V/ 


7^~ 












D? 


^ 


<?: 


s^,,' 












S^ 


=s^ 


r^"' 


--* 





































1.0 



0.8 0.6 0.4 

Load under Throttle 



0.2 



I300 1 
1200' 
1100' 
1000' 



50 



40 



30.-? 



20 



Fig. 191.- 



-Influence of compression pressure on charge dilution at various air 
densities and loads. 



than at full load if a satisfactory explosion is to be obtained. It 
seems probable that the air-fuel ratio for maximum economy does 
not fall below 15 for any operating condition that is likely to be 
met; that is, the maximum-economy mixture approximates to 
the maximum-power mixture as the air density and load decrease. 
With this in mind the performance curve for carburetors under 
partial loads can be examined. It would appear that constancy 
of mixture ratio under varying load is not desirable, and that 
1 P. S. Tice, loc. cit., p. 634. 



264 



THE AIRPLANE ENGINE 



a carburetor should show enrichment of the mixture with 
diminishing load. 

Performance of Representative Carburetors. — Several carbu- 
retors have been investigated at the Bureau of Standards 1 to 

20 



16 



14 



£12 
JlO 

























































% 

^s^ 


^ 
















• 




<S« 


























V>A 


^ 


Load 

1.0 

0.8 

0.6 

0.4 








• ^ 








• ^ 


























0.2 
0.1 



0.08 0.07 0.06 0.05 0.04 0.02 

Air Density in Pounds per Cu. F+. 

Fig. 192. — Variation of air-fuel ratio in Zenith carburetor. 

ascertain the variation in air-fuel ratio with variation (1) of 
air density and (2) of load. Three of these carburetors are 
considered here. The Zenith carburetor, A, Fig. 177, which is 

20 



\16 

< 

"14 
o 



8 



1 












o= 


0.O7^n 














m 




= 00702* 












I 








D=006I4 














1 














D= 0.0520 


















1 ~ 


( 








£z2£*es 


























5^ 


2327 










> 















































1.0 



0.6 0.4 

Load under ThroHle 



0.2 






Fig. 193. — Variation of air-fuel ratio in Zenith carburetor. 

described in detail on page 272, has two jets, of which one is 
operating under constant discharge head to compensate for the 
natural enrichment of the mixture with increase of load which 
would take place if the other or main jet alone were used. The 
1 P. S. Tice, loc. cit, pp. 620-636. 



THE CARBURETOR 



265 



Stewart- Warner carburetor, B, Fig. 177, has the throttle in the 
intake (anterior) and compensates for load changes by reducing 



26 



26 



15 24 



.b22 
< 



o20 



16 



12 



10 

























1.0 












































0.95 


















Load 
1.0 






































0.2 




0.8 
0.6 


















0.8 




0.4 
0.2 


















0.4 






















0.6 



























0.08 0.07 0.06 0.05 0.04 0.03 

Air Density in Pounds per Cu. Ft. 

Fig. 194. — Variation of air-fuel ratio in Stewart- Warner carburetor. 

the air pressure in the float chamber as the load increases by 
means of a passage connecting the choke discharge to the float 
chamber. The Stromberg carburetor, C, is described in detail 

28 

26 

24 
U 

^22 

i- . 

I 20 
o 

"§18 

cu 

X 

'£ 

14 
12 
10 



























1 




























































































§k 


&0t. 


?*-, 


















S^fcS 


















"^ 


^D=t 


10350 

































1.0 0.8 0.6 0.4 0.2 

Load under Thro+fle 

Fig. 195. — Variation of air-fuel ratio in Stewart- Warner carburetor. 



on page 278. The results of the investigations are exhibited in 
Figs. 192 to 197. For each carburetor there is shown the varia- 



266 



THE AIRPLANE ENGINE 



tion of air-fuel ratio with constant throttle opening and variable 
air density, and with constant air density and variable throttle 
opening. The absolute values of the air-fuel ratio are unimpor- 



16 



>' 4 

< 

ii I 2 
o 



? 8 
x 

























































































Load 






















1.0 
0.8 






















0.6 
0.4 

























T 0.08 



0.07 . 0.06 0.05 0.04 0.03 

Air Densi+y in Pounds per Cu. Ft. 



Fig. 196. — Variation of air-fuel ratio in Stromberg carburetor. 

tant in this connection since they are controlled by the size of the 
fuel jet, which can be readily changed; the method of variation 
of that ratio may, however, be considered as characteristic of 
each type of carburetor. In the Zenith and Stromberg carbu- 



16 



>I4 

< 



c?!0 



















































































>D =0.0758 




















0= 0.0498 


f D=0.0 


704 
















D = 0.030/ 





























1.0 0.8 0.6 0.4 0.2 

Load under Throf-He 

Fig. 197. — Variation of air-fuel ratio in Stromberg carburetor. 



retors, the need for an additional altitude control device is 
obvious; the mixture ratio at full load varies from 19 to 10.5 in 
the Zenith (Fig. 192) and from 15.5 to 9.5 in the Stromberg 



THE CARBURETOR 267 

(Fig. 196) as the air density diminishes from 0.07 to 0.03. The 
enrichment is considerably in excess of that which has been shown 
(Fig. 188) to be necessary. With load variation at constant air 
density, the mixture is practically constant in the Zenith car- 
buretor (Fig. 193), but enriches with diminution of load in the 
other two (Figs. 195 and 197) ; it has previously been shown (p. 
263) that such enrichment is desirable. 

Altimetric Compensation. — The importance of maintaining 
an economical mixture at high altitudes is attested by general 
experience in the air. British tests, to ascertain the advantages 
of a special altimetric control of the carburetor, have shown with 
water-cooled engines an increase in endurance from 4 to 4% hr., 
and in ceiling from 19,000 to 21,000 ft.; with air-cooled cylinders 
an increase in endurance from 5J-2 to 6% hr., of ceiling from 
15,000 to 18,000 ft. and of speed from 84 to 92 miles per hour. 
In addition to this there is less fouling of the spark plugs, the 
cylinders keep cleaner, and there is less danger of stalling the 
engine. 

Viscous Flow Carburetor. — It has been shown (p. 259) that 
with a standard simple carburetor with sharp-edged fuel orifice, 
the air-fuel ratio varies as the square root of the air density with 
full throttle and constant engine speed. There is a possibility 
of making this ratio constant, under varying air density, by sub- 
stituting for the sharp-edged orifice a capillary passage in which 
the flow is entirely viscous. The laws of viscous flow are com- 
plicated, 1 but, with velocities below those of turbulent flow, it is 
approximately true that the velocity of flow is proportional to 
the pressure head. In that case, referring to page 259, we have 

W x _ D* _ w x 
Wo D Wo 

W x Wo 

or — - = — -, that is, the air-fuel ratio remains constant with 
w x w ' ' 

varying air density. 

Carburetors have been built embodying the above principle, 
the viscous flow being obtained by the use of long capillary tubes, 
or by flow between flat discs or cones as in Fig. 198. At partial 
loads the fuel supply will fall off in proportion to the decrease in 
pressure head instead of in proportion to the square root of the 
pressure head and the mixture will consequently be too weak at 
low loads. Load control is obtained by raising or lowering the 

1 See Herschel, Bureau of Mines, Technologic Paper 100. 



268 



THE AIRPLANE ENGINE 



disc (or cone) of Fig. 198 and thereby changing the width of the 
capillary passage; this can be done by interconnection with the 
throttle lever. The principal objection to this type of carburetor 
is that the fuel flow varies with the fluidity of the oil and this 
varies both with the grade of oil used and with its temperature 
(Fig. 186). A further difficulty is sluggishness in response to 
quick opening or closing of the throttle valve. 





Fig. 198. — Diagrams of viscous flow carburetors. 



Altimetric Control. — The only practical method at present 
available for adjusting the air-fuel ratio to the desired value 
at all air densities, as well as at all throttle positions, is by the 
use of an additional or altimetric control. A carburetor may be 
designed so as to give correct mixtures for varying load or for 
varying air density but it cannot satisfactorily meet both con- 
ditions, since, with the same weight of air flowing the weight 
of fuel will be different in the two cases. For example, the 
weight flow of air at half load at the ground will be the same as 
at full load at an altitude where the air has half ground density; 
the pressure drop and the fuel flow will, however, be different 
in the two cases and therefore the air-fuel ratio will be different. 
If the carburetor is designed to give correct mixture at all alti- 
tudes at full load there would have to be added to it a load con- 
trol (preferably connected with the throttle valve) which would 
enrich the mixture at partial loads. The other method of pro- 
cedure is, however, usual; the carburetor is designed to give cor- 
rect mixtures at full and partial loads, and an altitude control 
is installed to permit a diminution in the fuel supply at higher 
altitudes. This control is nearly always manually operated 
but it can be made automatic without much complication. 

A diminution of fuel supply can be brought about either 



THE CARBURETOR 



269 



(1) by diminishing the size of the fuel orifice, or (2) by controlling 
the pressure head under which the fuel is flowing. The former 
is most readily accomplished by the use of a needle in the jet; 
the latter is the method generally employed because it is less 
sensitive in adjustment and turns out to be more robust as a 
structure. 




Fig. 199. 



A B 

-Altitude control by regulation of the float-chamber pressure. 



Schematic diagrams of some of the more promising methods 
of altitude control are shown in Figs. 199 and 200. x For the 
control of the float-chamber pressure, Fig. 199, the top of the 
float chamber must be provided with a vent, a, to the atmosphere, 
and a connection, b, to some place where the pressure is less 
than atmospheric. The control valve may be in either of these 
passages. 





ABC 
Fig. 200. — Altitude control by regulation of the jet discharge pressure. 



The nozzle outlet pressure can be controlled in several ways. 
The position of the outlet relative to the air passage can be 
changed, either by shifting the choke, as in Fig. 200A, or 
by shifting the outlet. The amount of air passing the outlet can 
be reduced by the use of an auxiliary air valve located at a point 



National Advisory Committee for Aeronautics, 4th annual report, p. 637. 



270 THE AIRPLANE ENGINE 

beyond the fuel outlet, as in C. A third method is to admit 
(or bleed) air to the fuel jet past the metering orifice, as in B, 
thereby reducing the pressure head on the orifice. 

The structures involving a small plug valve controlling an 
air stream (Figs. 199 and 2005) are the simplest and most easily 
produced. Their regulation is comparatively direct and involves 
small forces and a minimum of parts; furthermore, they adapt 
themselves readily to automatic control. For such reasons, these 
methods are the ones usually encountered in service. The 
objection to them is that they do not permit of one setting for all 
loads at any given air density but require adjustment for each 
throttle position, if maximum economy is to be maintained. 

The method of Fig. 200A is structurally clumsy and would 
complicate the carburetor considerably. The method of Fig. 
200C, using a balanced auxiliary valve, would offer little resistance 
to operation and little complication. Moreover, the mixture 
should be satisfactory at partial loads without further manipula- 
tion. The auxiliary valve would have to be large to give com- 
plete compensation up to one-half ground density. A simple 
calculation shows that for this range the area of the auxiliary 
port must be approximately 1.5 times that of the carburetor 
throat. 

Manual operation of the altitude control is extremely unde- 
sirable. The operation should be continuous as the plane 
changes its altitude or speed and can at best be only intermittent 
with manual operations. Moreover, the pilot has no definite 
means of knowing how far to move the control but must rely 
chiefly on the engine tachometer readings. He can find the 
maximum power position but not the more important maximum 
economy position. As he is already burdened with a large num- 
ber of controls it is much better to make the altimetric compensa- 
tion automatic. 

The simplest automatic operating device is an aneroid bellows. 
A sealed flexible-walled chamber will expand under reduced 
pressure and under increased temperature, that is, it will respond 
to change in air density. If correction for pressure only is 
desired, the bellows can contain a spring under compression and 
can be exhausted before sealing (see Fig. 212). Such devices can 
only operate satisfactorily if the resistance which they have to 
overcome is small and if the method of control is such as not to 
disturb the compensation at partial loads. 



THE CARBURETOR 271 

Atomization. — The preceding discussion has concerned itself 
with the metering or mixture-making characteristics of carbure- 
tors. Other qualities which are of importance are (a) the degree 
of atomization of the fuel and the homogeneity of the mixture; 
(6) the pressure drop through the carburetor at wide-open 
throttle; (c) satisfactory idling performance; (d) acceleration. 
All carburetors, in order to be acceptable, must be satisfactory 
not only in mixture making but also in these other characteristics. 
Favorable conditions for fine atomization of the fuel are high 
velocities of the air and, to a minor degree, of the fuel. The air 
velocity is always much greater than that of the entering fuel and 
the atomization is largely due to the high relative velocity of the 
air. This is particularly marked if the fuel is not discharged in 
the axial direction. The use of an anterior throttle, as in Fig. 
1775, by increasing the air velocity at the jet improves atomiza- 
tion at partial loads. The admission of air before the fuel outlet 
but past the orifice (see Fig. 200B) is a further favorable condition. 

Good atomization may be impaired by the impinging of the 
mixture on obstacles such as a butterfly throttle valve, placed 
centrally above the jet (see Fig. 177A). Here again an anterior 
throttle has an advantage. The mixture will impinge on the 
inlet manifold and the valves before getting into the cylinder, but 
it is better to have such actions take place as far away from the 
mixing point as possible. Best results have been obtained with 
a long pipe leading from the carburetor to the manifold, giving 
more time for vaporization and the formation of a homogeneous 
mixture before the mixture is taken into one or other branch of 
the manifold. 

Pressure drop through the carburetor has been touched on in 
page 251 in the discussion of the discharge characteristics of the 
air passage. Its importance is solely in affecting the maximum 
power output. 

Idling. — An engine requires a richer mixture at lighter loads. 
When the engine is cold a still richer mixture is necessary. None 
of the carburetors in use on airplanes will give a satisfactory idling 
mixture without the use of some auxiliary device. This consists of 
a fuel discharge above the throttle which utilizes the high vacuum 
above the closed throttle to suck in the necessary amount of fuel. 

Acceleration. — It is of importance that the mixture should 
respond rapidly to sudden changes in load. If the throttle valve 
is opened suddenly, the greater density and inertia of the fuel 



272 THE AIRPLANE ENGINE 

tend to make the mixture too weak, with the result that the 
engine will back-fire or misfire. To avoid this, it is common to 
have an auxiliary supply of gasoline which, at partial loads, 
collects near the fuel outlet and is drawn on first when the 
throttle is suddenly opened, keeping up the strength of mixture 
until the regular flow is established. 

Certain special conditions have to be met with by an airplane 
carburetor as a result of manoeuvres of the plane. The changing 
inclination of the plane will change the hydraulic head at the jet 
unless it is placed at the center of the float chamber. With the 
usual non-concentric arrangement of parts (see Fig. 177) it is 
desirable to have the float chamber placed in advance of the jet 
as this will give a greater hydraulic head and richer mixture on 
climbing and will cut down the fuel supply on descent or diving. 
It is necessary to see that the gasoline does not overflow from 
the jet when the plane is resting on the ground. The action of 
the float and float valves during a dive must be examined. The 
usual float, guided by a central spindle which is normally vertical, 
will go out of action during a dive, with the probable result of 
flooding the carburetor. Special float mechanisms are desirable 
and have been devised. In case of flooding during a dive, the 
air horn or intake pipe should be so arranged that gasoline can- 
not spill out into the fuselage. As the air horn is usually facing 
forward to get the advantage of the increased air pressure due to 
the relative wind velocity, such spilling will occur unless the air 
intake pipe is led upward before being turned forward. 

The usual dual carburetor has one float chamber, and one 
air intake to the two chokes. A dual air intake pipe is to be 
recommended as reducing the risk from back-fire, by making each 
group of three or four cylinders a separate unit so far as carburiza- 
tion is concerned. With a common air pipe, back-fire may cause 
the engine to stop; with double intake, back-fire into one intake 
will not interfere with the operation of the cylinders fed from the 
other intake, the engine continues to run and the flame in the 
back-firing intake is drawn up into the engine, reducing the risk 
of fire. Furthermore, a dual intake increases engine power by 
diminishing the resistance to air flow. 

CARBURETOR CONSTRUCTION 

Zenith. — The carburetor which has been used most for air- 
plane engines is made by the Zenith Carburetor Co. In this 



THE CARBURETOR 



273 



carburetor, an attempt is made to maintain constant mixture 
strength at varying throttle positions by the use of two jets or 
nozzles, one of which, the main jet, acts in the usual way, while 
the other, the compensating jet, delivers an amount of fuel which 
is entirely independent of engine speed and load. This arrange- 
ment was devised by Baverey in 1906. The main jet alone would 
give a mixture which is at all times too weak, but which becomes 
richer as the engine speed and load increase; the compensating 
jet alone would give a mixture which is at all times too weak but 
which becomes weaker still as the engine speed and load increase. 
The two jets working together tend to compensate one another, 
and, if properly proportioned, will give a mixture of fairly constant 
strength under varying speed and load. This is shown in Fig. 
193. In this case, the jet sizes are No. 140 for the main jet and 





(a) (b) 

Fig. 201. — Diagram showing action of the Zenith carburetor. 



No. 150 for the compensating jet, the number indicating the 
cubic centimeters of water discharged per minute under a 12-in. 
head. The discharge for the compensating jet is under a con- 
stant head of 2 or 3 in. of water; the main jet discharge is under 
the variable head due to the pressure drop at the throat of the 
venturi, which depends on the size of the throat and may amount 
to 40 in. of water in usual designs. The arrangement of these 
jets is shown diagrammatically in Fig. 201, in which a shows con- 
ditions at rest, and b at full throttle. The main jet, G, is located 
as usual; the compensating jet, I, discharges into the well, J, which 
empties into a nozzle, H, concentric with the main jet, G. When 
at rest, the levels in the float chamber, the wells, and the nozzles 
G and H, are the same. On opening the throttle, the capacity of 
the nozzle, H, is so much greater than that of the jet, I, that the 
well, J, is kept drained and both air and fuel are sucked up the 

18 



274 



THE AIRPLANE ENGINE 



nozzle, H. As the pressure in the well, J, is atmospheric, the dis- 
charge through I is due to the hydrostatic head of the liquid in 
the float chamber and is therefore constant. The well, J, serves 
also as an accelerating well, giving a body of fuel immediately 
available on opening the throttle from the idling position. 
At low speed, when the throttle valve, T, is nearly closed, the 
suction at the throat is not sufficient to draw in any gasoline 
and it enters only through the idling device. This device, shown 
diagrammatically in Fig. 202a, consists of the idling tube, M, 
within the secondary well, P, which is inserted in the main well, J, 
into which the discharge from the compensating jet, 7, occurs. 
The well P is provided with a small metering orifice at the bottom 
through which gasoline can enter from J, and with small air 





M 



(b) 



Fig. 202.- 



-Diagram showing (a) idling device and (b) altitude control of the 
Zenith carburetor. 



holes at the top. The idling tube, M, terminating opposite the 
throttle valve, is subjected to a very strong suction whenever the 
throttle is nearly closed and discharges gasoline from the well P. 
This gasoline meets the air passing with great velocity through 
the small opening around the throttle valve and forms the idling 
mixture. As the throttle is opened, the vacuum at the throttle 
diminishes while that in the choke increases, so that discharge 
through M ceases and that through G begins. 

The altitude control of the Zenith carburetor is shown diagram- 
matically in Fig. 2026. It is of the type illustrated in Fig. 199A. 
The float chamber is open to the air through screened air inlets. 
The well J is in open communication at its top with the float 
chamber. A passage, P, from the float chamber to the choke 
discharge, is fitted with a stop cock, L, which is manually operated 
by the pilot. This cock is closed at the ground and is opened 



THE CARBURETOR 



275 



gradually as higher altitudes are reached; it should be opened as 
far as is possible without appreciably diminishing the revolutions 
of the engine. 

The actual construction of a Zenith carburetor is shown in 
Fig. 203. Gasoline enters the float chamber through D and the 
needle valve seat, S. As soon as it reaches a predetermined 
height the metal float, F, acting through the levers, B, and the 
collar, iV, closes the needle valve, C, on its seat S. From the float 
chamber the gasoline flows (1) through the compensating jet, I, 




Fig. 203. — Section of Zenith carburetor. 

into the bottom of the well, J, and then through the channel, K, 
to the cap jet, H, which surrounds the main jet, G, and (2) through 
the channel, E, to the main jet, G. The idling tube, M , is inside 
the secondary well,P, and discharges through the passage, R, to an 
opening (not shown) opposite the throttle valve. The altitude 
control valve, F, is a tube which is shown communicating with 
the choke discharge; the other communication to the float 
chamber is not shown. It is operated by the lever X. 

As in other carburetors, a single float chamber is used to supply 
two air chokes if the engine has six or eight cylinders. One air horn 



276 THE AIRPLANE ENGINE 

or intake commonly serves the two chokes of a duplex carburetor, 
but it has been found that greater engine power can be obtained 
if separate intakes are used. Tests of special Zenith carburetors 
for the Liberty engine showed maximum power developed with 
separate air intakes about 4 in. long. 1 

The special feature of the Zenith carburetor which has recom- 
mended it is the absence of all moving parts. It is general 
experience that auxiliary air valves, metering pins, and other 
moving devices will stick at times and cause irregularity of 
action. For maximum reliability and fool-proofness the com- 
pensating device should be fixed. 

The Claudel carburetor, which has been used very extensively 
for airplane engines, especially in Europe, is now being made in 
this country. Like the Zenith, the compensation for load and 
speed is made without any moving parts. A general view is 
shown in Fig. 204. The fuel discharges into the choke from a 
diffusor which is shown assembled in Fig. 2056. The diffusor 
has four concentric tubes, the air tube e, guard tube d, diffusor 
tube c, and idling tube a. The main jet is in a small plug screwed 
into the bottom of the diffusor. Air at atmospheric pressure 
enters the bottom of the air tube, passes over the top of the 
guard tube (which prevents the fuel from overflowing when the 
engine is at rest), then goes through such holes in the diffusor as 
are above the fuel level, and out through the nozzle holes to the 
throat of the venturi. The fuel is at the level shown when the 
engine is idling or at rest. As the throttle is opened, the suction 
in the diffusor increases, thereby lowering the liquid level in the 
diffusor bore and uncovering progressively a series of air-bleed 
or compensating holes. Through these holes the air rushes into 
the ascending column of fuel and atomizes it as it leaves the 
nozzle holes at the top. At maximum load the diffusor is practical- 
ly emptied and all the air-bleed holes are in action, cutting down 
the effective head on the fuel. The compensation is by control- 
ling the jet outlet pressure along the lines indicated in Fig. 
2005. Any desired kind of compensation can be obtained by 
appropriate ^design of the size and location of the air-bleed 
holes. 

The diffusor acts also as an accelerating well. When idling 
the diffusor is out of action and all the fuel goes through the cen- 

1 Bulletin, Experimental Department, Airplane Engineering Division, 
U. S. A., Jan., 1919. 



THE CARBURETOR 



277 



tral idling tube, mixed with some air entering compensating holes 
from the air tube. 




YMUtffrmrt 



Fig. 204. — Section of Claudercarburetor. 



Idling Tube . 

Compensatfngfyi 
L- Hofes 



Idling 
Jet. 

Main 
'"Jet 



Nojj/e Hales Compensating Males 



(a) 



I Fuel level 



Bore 



Guard 
Tube 







Air Tube 



TJ 



Diffusor 
Main Jet Guard Tube PI uq, 

(b) * (c) (d) (e) 

Fig. 205. — Details of Claudel diffusor. 



The throttle is a cylindrical or barrel throttle, bored out so as to 
form a smooth continuation of the venturi when it is wide open. 
It offers no resistance at maximum load and consequently leads 



278 



THE AIRPLANE ENGINE 



to maximum volumetric efficiency and power. As the idling 
tube projects into the throttle space, the throttle is slotted out 
wide enough to pass around it. To diminish the area through 
this slot when the engine is idling a screw, c, extends into the air 
space. Advancing the screw lessens the air area and enriches 
the idling mixture. Figure 206 shows the idling position. 

Another feature of this carburetor is the sliding air cone, A 
(Fig. 204) , which is controlled by an external lever. When the 
cone is raised to contact with the venturi, it shuts off all air 
supply and puts maximum suction on the diffusor. This greatly 
enriches the mixture and is advantageous for starting in cold 





Fig. 206.- 



-Idling device of the Claudel 
carburetor. 



Fig. 



207. — Section of dual Claudel 
carburetor. 



weather. The same device is used for altitude control. The 
venturi used in airplanes is larger than is necessary at the 
ground. At low elevations the air cone is kept in a raised 
position in order to increase the suction in the diffusor to the 
amount necessary to give the desired mixture. As elevation is 
gained the air cone is gradually lowered, thus compensating for 
the natural increase in richness. 

A cross-section through the diffusors and throttle valves of a 
duplex Claudel carburetor as used on the Hispano-Suiza engine 
is shown in Fig. 207. 

The Stromberg carburetor, Fig. 208, although structurally 
very different, uses the same general method of compensation 



THE CARBURETOR 



279 



for speed and load as the Claudel. The special features of this 
carburetor are the float mechanism and the double venturi. 

The float (Fig. 208) is spherical or cylindrical (with horizontal 
axis) and is hinged as shown with the pivot toward the tail of the 

" Idle Tube Abjj/e 
Discharge 
Nojj/e 
Large Vznturi 
Tube — ... 




\AcdleraHn 3 Mai Z M ' i * nn 3 
Metering No JJ /e 

Air Horn Drain Mojj/e 

Fig. 208. — Section of Stromberg carburetor. 

plane. With this mounting, the float is in action during all 
ordinary manoeuvres of the plane (Fig. 209), that is, it keeps 
the needle valve closed with a moderate amount of gasoline in 




Fig. 209. — Diagram showing Stromberg float chamber in different orientations. 

the chamber. If the plane goes upside down the weight of the 
float will close the valve. With the arrangement of Fig. 208 
the main jet will overflow into the air inlet during a steep dive 
with closed throttle. A duplex carburetor arranged as in Fig. 



280 



THE AIRPLANE ENGINE 



210, with the float between the two discharge jets, leaves no 
possibility of such leakage of fuel. 




Large venturi 
tube 



Small venfuri'' 
tube 



Accelerating well- 
Metering nozzle- 



Altitude control tube' 
tlain gasoline channel 



'Float 

Air horn drain connection 



Fig. 210. — Section of dual Stromberg carburetor. 



The diagrammatic sketch (Fig. 211) shows the metering jet, 
E, discharging into channel, A, with air-bleed holes, D, through 
which air at atmospheric pressure enters from the outer channel, 
B. The outer channel is also the accelerating 
well. 

The fuel and the atomizing air are dis- 
charged radially into the choke through a ring 
of small holes, located at the throat of a small 
venturi tube. This small venturi is concen- 
tric with a larger venturi and discharges at 
its throat. The discharge pressure of the 
small venturi is considerably below atmos- 
pheric pressure and the depression is still 
greater at the throat of the small venturi. 
This results in very high velocity for that por- 
tion of the air supply which passes through 
the small venturi, giving good atomization of 
the fuel without having to make the whole 
air supply acquire a very high velocity. This 
arrangement gives a small total pressure drop in the carburetor, 
and consequently high volumetric efficiency of the engine. 

The idling device is a miniature carburetor with discharge 
just above the closed throttle. The idling tube connects directly 




Fig. 211. — Dia- 
gram showing load 
control of Stromberg 
carburetor. 



THE CARBURETOR 



281 



with the main jet passage and has a fuel nozzle discharging into 
a mixing chamber where it meets air entering through holes 
which are controlled by a needle valve. The discharge nozzle 
into the main choke is a slot of which more is exposed as the 
throttle moves from its closed position. The increased opening 
of the slot increases the suction in the mixing chamber, and 
sucks up more fuel as the throttle begins to open. With still 
further opening the suction at the main discharge nozzle increases 
while that at the idling nozzle decreases. There is a throttle 
position at which fuel discharges through both, but with still 
further opening the idling nozzle goes out of action. 




A 3 j 



Fig. 212. — Diagram showing automatic altitude control attached to Stromberg 

carburetor. 



Altitude compensation is effected by controlling the pressure 
in the float chamber. An arrangement for automatic control 
is shown diagrammatically in Fig. 212. The aneroid chamber, 
A, which has been exhausted before sealing, is compressed by 
the joint action of the air pressure and the spring B. As the air 
pressure diminishes the aneroid expands compressing the spring 
and raising the valve C. The valve point is slotted and offers a 
decreasing aperture for the admission of air as the valve rises. 
Air is sucked through this slot by the action of the venturi at D, 
and as the only air vent from the float chamber is into the pipe E, 
the pressure in the float chamber will vary with position of the 
valve C. An additional manual control is a necessary safety 
device. 

The design in Fig. 210 is especially adapted to a 90-deg. Vee 
engine and, as previously pointed out, permits a position of the 



282 



THE AIRPLANE ENGINE 



float chamber between the two carburetor outlets which largely 
eliminates the disturbing factor of changing inclinations of the 
plane. The carburetor barrels are water-jacketed for high 
altitude service. The main fuel nozzles are in an annular groove 
around the small venturi. The altitude-control suction is 
through the small axial tubes shown terminating at the throats 
of the small Venturis and consequently give the maximum possible 
suction and range of action of the control. The altitude control 
has a partial connection with the throttle in such way that the 



l 


~ , 




J u 


f 


1 s 




r 


H 




s — '-" 


















>\ fig 




Fig. 213. — Sections of Miller carburetor. 



mixture is enriched during the latter part of the closing of the 
throttle. 

The Miller carburetor has been used on the U. S. Bugatti 
engine. It is of the multiple-jet type in which load compensation 
is effected by bringing more jets into action as the throttle is 
opened, the sizes of the jets being designed to give correct mix- 
ture at all loads. The jets are air-bled, giving compensation 
for varying speed. The jets are held in a narrow holder (Fig. 
213) and discharge across a diameter at the throat of the venturi. 
The drill sizes for the Bugatti engine are No. 76, which is the 
idling jet, No. 76, No. 75, No. 71, No. 68, No. 57, No. 53. The 
corresponding diameters in inches are 0.020, 0.021, 0.026, 0.031, 



THE CARBURETOR 



283 



0.043, 0.0595; the areas consequently increase very rapidly. 
These jets come into action progressively as the throttle is 
opened. Each jet has four small air holes just above the metering 
orifice; air enters at atmospheric pressure through a %Q-m. 
hole near the top of the jet holder and passes down around the 
outside of each jet to the air holes. The gasoline flows from the float 
chamber to the lower %6 _m - hole in the jet holder. The idling 
jet is the first in the holder. 

The throttle valve is of the barrel type bored out to give a 
venturi form when wide open. The stop for the idling position 
is seen in the figure. Altitude compensation is obtained by 
varying the pressure in the float chamber, the air space of which 
is at all times in direct connection with the venturi. A manually- 




Floaf- 



Air Damper 
Gas Passage Air Make 

Fig. 214. — Sections of Master carburetor. 



operated valve controls the size of the free air connection to the 
top of the float chamber. 

The Master carburetor is also of the multiple-jet type, but 
differs from the Miller in that the jets are all of the same size 
and are not air-bled. The throttle is of barrel type (Fig. 214) 
with an opening that is curved so as to uncover the jets pro- 
gressively as the throttle is opened. An air damper controlled 
by the pilot restricts the venturi opening and consequently 
enriches the mixture when desired for starting. The number 
of jets is usually from 14 to 21, which demands extremely small 
metering orifices. 

The Ball and Ball carburetor (Penberthy Injector Co.) is of 
the single metering orifice, air-bled type. The float is spherical 
in a spherical chamber. The venturi throat, A, (Fig. 215), has 
the main nozzle tubes, B, connecting through the annulus, C, 



284 



THE AIRPLANE ENGINE 



with the passage, D, and the mixing chamber, E. The metering 
jet, F } is at the bottom of the nozzle, G, and the fuel overflows 
through the four air holes, H, into the chamber, E, which connects 
to the outside air through the passage, M, and the air orifice, N. 
Gasoline arrives from the float chamber at J. The idling jet, P, 
connects through the passage, 0, with the mixing chamber, E, and 
discharges just above the closed throttle. An auxiliary air 
valve, S, is sometimes used to reduce the strength of the mixture 
at heavy loads. 




Fig. 215. — Section of Ball and Ball carburetor. 



The altitude control is by variation of the pressure on the dis- 
charge side of the main jet. This is accomplished by substitut- 
ing a larger valve-controlled opening for the air orifice, N ; opening 
this valve increases the pressure on the discharge side of the main 
jet and weakens the mixture. 

The carburetor used on the (German) Basse-Selve engine is 
simpler and lighter than any of the types previously discussed. 
The float (Fig. 216) is annular, and concentric with the choke, 
thereby reducing the possibility of overflow of gasoline from the 
main jet when the carburetor is inclined. The main jet is 



THE CARBURETOR 



285 



formed by a hole drilled in a tube which is screwed diagonally 
into the water-jacketed body of the carburetor and lies across 
the choke tube. The jet tube is open at its lower end and 
projects into the bottom of the float chamber. The idling jet 
is formed by a second tube of small diameter inside the jet tube. 
This idling tube is also open at the bottom and is drilled radially 
with a small hole just below the main jet. It communicates 
with the mixing chamber just above the throttle by a passage 
drilled in the carburetor body. Altitude compensation is by 
varying the air pressure in the float chamber. 




Fig. 216. — Sections of Basse-Selve carburetor. 



The float chamber is made of pressed sheet steel of very light 
gage. The needle valve (Fig. 216) is acted on directly by the 
float without the intervention of levers. 

The carburetor of the Bayerische Motoren Werke engine has 
some noteworthy features. It consists of three carburetors 
with a common float chamber (Fig. 217). Each of these car- 
buretors has a separate discharge pipe leading to a common 
induction manifold. The central carburetor has both idling and 
main jets; the outer two have main jets only. There are five 
throttle valves arranged in two systems with independent control. 
The main system has three throttles, one to each carburetor. 
The secondary system, which is an altitude control, has valves on 
the outer carburetor only. 

The action is as follows: When the main throttle is opened 
slightly, the side throttles remaining closed, the idling jet (center 
carburetor) alone is in action; mixture from the center carburetor 



286 



THE AIRPLANE ENGINE 



alone reaches the cylinders. As the throttle is opened further 
the main jet of the center carburetor comes in action and supplies 
the whole mixture until the throttle is half open. After this, 
the two side carburetors, which are controlled by slotted links, 
begin to open. The normal continuous ground level full-power 
operation is at the point where the side jets are just about to 
begin to discharge. 

So long as the secondary throttles remain in their closed 
position with relatively small passages past them, a compara- 
tively rich mixture is supplied by the side carburetors. As 
altitude is gained the secondary throttles are opened and give 
increased power while keeping the mixture of the desired strength. 



Adjuslmerrf- 

ibr Slow Funning p 

Pi lot Jet -. 




MamJe-hs 



Fig. 217. — Sections of B.M.W. carburetor. 



An entirely different type of carburetor is used on the Maybach 
engines on large German dirigibles. These have been designed 
to dispense with the use of a float chamber and to work in con- 
junction with a gasoline-pump system. The construction is 
shown diagrammatically in Fig. 218. The throttle valve, /, is of 
the rotary-barrel type and admits carbureted air from N and 
fresh air from L. The throttle lever is interconnected with the 
sliding shutter, K, controlling the air that flows past the jets, and 
with a rotatable cover, P, regulating the size of the jets. Fuel 
from the gasoline pump enters an upper vessel, A, by the pipe, B. 
The level in this vessel is kept constant by an overflow pipe, C, 
which conducts the excess fuel back to the supply tank. An air 
vent fitted with a baffle plate is provided at F. The fuel passes 



THE CARBURETOR 



287 



from A through a strainer, M, to the vessel, D, whence it is 
sucked through the orifice, F, into the induction pipe. Excess 
fuel in D overflows and joins the excess from A in the pipe C. 
At the top of vessel D two holes are drilled — the main and idling 
jets. These orifices are controlled by the eccentrically-mounted 
cap, P, which is rotated through interconnection with the throttle 




Fig. 218. — Diagram of Maybach carburetor. 

lever. The fuel has a constant liquid head equal to the difference 
in levels between the liquid in A and the level of the orifices; in 
addition it is subjected to the suction in the passage above H. 

In the idling position, L is open slightly (Fig. 219), K is closed, 
and the idling jet only is uncovered by P. The throttle-lever 




Angular displacement of control lever. 

Fig. 219. — Action of the Maybach carburetor. 

quadrant is marked with the positions "idling," "low speed," 
"full power," and " altitude." As the throttle is rotated from 
the idling position, which demands a rich mixture, the shutter K 
opens but the fuel opening does not increase much till the "low 
speed" position is reached; the fuel discharge increases in con- 
sequence both of increased fuel orifice and of the increased 



288 



THE AIRPLANE ENGINE 



suction at H. The "full power" position is not maximum power 
but is the maximum at which it is desirable to operate the engine 
at ground level. The fuel orifice is nearly wide open at the full- 
power position. With further opening of the throttle the fresh- 
air inlet L opens more, thereby preventing the enrichment of the 
mixture which otherwise would occur at high altitudes and 
maximum power. 

It is evidently possible to design the dimensions and the 

interconnections of the three 
orifices G, N and H in such 
way as to give any desired 
mixture to an engine operat- 
ing at ground level and at 
maximum power at various 
altitudes. Partial loads at 
high levels are not provided 
for. This method of meeting 
the carburetor problem is un- 
desirable because of the com- 
plexity of the design and the 
practical impossibility of mak- 
ing the varying fuel orifices of 
the desired dimensions. This 
particular carburetor is very 
heavy and offers a large air 
resistance, thereby reducing 
the volumetric efficiency and 
power of the engine which it 
supplies. 

A very simple type of carburetor is used on the rotary Le Rhone 
engine. The air-fuel mixture enters the rotating crankcase 
through a stationary hollow crankshaft. The screened air supply 
is controlled by a throttle which is in the form of a shutter (Fig. 
220) carrying at its lower end a long metering pin which controls 
the size of the fuel jet. The pressure at which the fuel arrives at 
the orifice is controlled by a by-pass valve; this serves to con- 
trol the mixture when altitude or load is changed. The inherent 
mixture control is irregular and uneconomical with a device of 
this nature. 




Fig. 220. — Section of LeRhone car- 
buretor. 



CHAPTER XI 
FUEL SYSTEMS 

The following statement of the requirements of the fuel 
system of an airplane engine is abstracted from the " Handbook 
of Instructions for Airplane Designers" prepared by the 
Engineering Division of the U. S. Air Service. 

There should always be more than one means of supplying 
fuel to the engine. 

Main -feed System. — Gravity feed should be used throughout if it is 
possible to maintain a sufficient head with the airplane at maximum angles 
of flight. It has been found that a head of 18 to 30 in. is required for 
satisfactory operation of current types of carburetor. Unless it is possible 
to maintain a sufficient head by gravity, pumps must be installed to supply 
gasoline from the main tanks to the engines. Pressure in supply tanks is not 
permitted on fighting planes. 

The main fuel pumps should have a capacity at least 50 per cent greater 
than the maximum requirement of the engines. Two pumps, other than 
hand pumps, are desirable, either of which can supply sufficient fuel. They 
should have automatic pressure regulation to eliminate the use of relief 
valves or other means of adjusting the pressure at the carburetor. The 
gasoline pressure at the carburetor must always be at least 1 lb. and the 
system should be so adjusted that this pressure can never rise above 3 or 
4 lb. as a result of change in position of the airplane. Pumps capable of a 
discharge pressure higher than 4 lb. should have relief valves connected 
between the suction and discharge, so adjusted as to limit the maximum 
discharge pressure to 4 lb. The fluctuation of pressure at the carburetor, 
due to pulsations of the pump, should not be over 25 per cent. Where air 
pressure is used, the power air-pump must be capable of keeping a pressure 
of 2 lb. on the tanks at the ceiling of the airplane and both spring- and 
manually-controlled relief valves should be furnished, the former set to 
relieve at 4 lb. per square inch. 

Pumps should preferably be located below the lowest point in the supply 
system. If they are located higher than the bottom of the main tanks, 
means must be provided for admitting gasoline from the auxiliary supply to 
the suction side of the pumps. It should be possible for the pilot to make use 
of this connection during flight. A non-return valve must be installed to 
prevent this gasoline from going into the main tanks instead of the pumps. 

Pumps which do not require glands are preferred, although satisfactory 
glands will be accepted; in case glands are used, they must be so located that 
any leakage from the glands will be drained to a point outside of the fuselage. 
Pumps should preferably be connected to and driven by the engine. 

Auxiliary-feed System. — The auxiliary-feed system supplies gasoline to 
the engine in case of failure of the main supply; this auxiliary system should 
19 289 



290 THE AIRPLANE ENGINE 

be such that fuel can be supplied to the engine in the shortest possible time 
never to exceed a period of 10 sec. from the time the pilot starts to make use 
of the auxiliary system. For emergency use, gravity tanks are best, but 
must not be used if a head of 12 in. in level flight is not obtainable; they 
should have sufficient capacity to operate the engines for 30 min. at an 
altitude of 10,000 ft. with wide-open throttle and should be so connected 
to the system that they can be shut off and used for reserve or emergency 
only. They must be so constructed or connected that they can be entirely 
emptied with the airplane inclined at maximum angles of flight. An over- 
flow pipe from the gravity tank returns any excess gasoline to one or more 
of the main tanks; this overflow must be so constructed that there can be 
no gasoline trapped in it when the airplane is in normal flying position. 

Unless there are three means of delivering fuel to the engine, such as two 
engine or wind driven pumps and a gravity tank, a hand gasoline pump must 
be provided which will permit the pilot, while controlling the airplane, to 
pump, without undue exertion, sufficient fuel from the main supply at proper 
pressure for the operation of all engines at full throttle. The capacity of this 
auxiliary system must be such that the pilot will not need to operate the 
pump during more than one-third of the time. 

Tanks should be of tinned steel and of such thickness that the tank will 
stand 5 lb. per square inch pressure on the inside without undue distortion. 
Flat surfaces are to be avoided. Wherever the width or length (horizon- 
tally) of a tank is greater than 12 in., a splash plate for reinforcing purposes 
must be installed at least every 12 in. ; wherever the height of a tank is greater 
than 18 in., a splash plate for reinforcing purposes must be installed at least 
every 18 in. All seams, including the connection between the splash plates 
and the walls of the tanks, should be riveted and soldered. Copper or soft 
iron rivets must be used throughout; the exposed parts of the rivets to be 
tinned in case iron rivets are used. 

Drains leading to a point outside the fuselage must be installed in the 
bottom of each main tank. Fillers must be conveniently located on each 
tank, and in such a position that the entire tank can be filled while the 
airplane is on the ground. A removable screen must be installed at the point 
of filling of each main tank, and also, if practicable, in the gravity tank. 
Vents must be located at the highest point on all tanks, usually in the filler 
tube, except on wing gravity tanks where the overflow pipe shall act as a 
vent. 

Line and Carburetor Strainers. — A line strainer, with removable screen 
and bowl, must be installed between the tanks and pumps, located as low as 
possible and in such a position as to be readily accessible for draining and 
cleaning. The strainer screen should be of brass, bronze or copper of about 
100-mesh and 0.005-in. diameter wire, and should have at least 1 sq. in. 
for each 6 gal. which must pass through per hour. 

Each carburetor should be provided with a strainer having a readily 
removable screen of brass, bronze or copper of about 50 mesh and approxi- 
mately 0.009-in. diameter wire and having an area of at least 2 sq. in. 

Service Pipes and Connections. — Service pipes should be % in. outside 
diameter where flow is 30 gal. per hour or less; x /l in. outside diameter where 
flow is between 30 and 60 gal. per hour; and % in. outside diameter where 



FUEL SYSTEMS 



291 



flow is between 60 and 100 gal. per hour. All vent and air tubes should be 
y± in. outside diameter. Wall thickness should be 0.028 in. for %-m., H 2 in. 
for ?£-in.. and % 4 in. for J^-in. and %-in. outside diameter. All service 
pipes, or tubing, should be seamless and of annealed copper, soft enough to 
withstand vibration. At all points where the tubing is connected to solidly 
mounted objects, such as pumps or tanks, flexible connections must be 
provided. The tubing must be properly protected at points of possible 
chafing. Sharp bends are not permitted. Tube fittings are to be of brass 
or bronze. 




Valve with 
spring to 
automatic- 
ally c/ose 

Distributing 
valve accesible 
top/ug 

Strainer 



Bottom of Fuselage 
Fig. 221. — Fuel system for a single-engine airplane with gravity feed. 



Multi-engine Installations. — When more than one engine is used, each 
should have its own gasoline system, consisting of pumps, main tanks, 
gravity or reserve tank, distributing valve and other apparatus required 
for a single engine system. A cross connection with shut-off valve should 
be provided so that any engine can take fuel from the tanks of the other 
engines, and unless two pumping units, not operated by hand, are provided 
for each engine, a cross connection should be provided so any engine may 
receive fuel from the pumps of the other engines. 

Priming Devices. — A priming system should be installed on every engine, 
with the priming pump mounted in the cockpit in an accessible posit'on. 



292 



THE AIRPLANE ENGINE 



Typical arrangements of the fuel system are shown in Figs. 
221 and 222. Figure 221 shows a system in which the carburetor 
is near the bottom of the fuselage so that gravity feed can be em- 
ployed. The auxiliary tank is most conveniently and simply 
made a portion of the main tank. A pump system with auxil- 
iary tank incorporated in one of the wings is shown in Fig. 222. 



Filler Cap-., Gravity Tank 



Jo 
Manifolds 




Pump priming va be to pilot 
this priming connection may 
be omitted if pvmps arebe/ow 

hnks and if the gravity tank 

atalltime can provide 

sufficient head at the 
carburetor 




Primer 



Overflowsightglass reacf//y 
visible topiiof 



Drain 





"' Valve with spring 
to automatically 
close 

Pressure Gage 



Distributing va/ye \ 
accessible to pi/of ) 



Lock Wire 
on cock handle 



.--J Way Cock 

2 Drain 
Bottom of fuselage 



Duplex Pump 
Fig. 222. — Fuel system for a single-engine airplane with pump feed. 

Pumps. — A simple form of air pump, used in the Hispano- 
Suiza engine, is shown in Fig. 223. It is operated by a cam 
on the camshaft which gives the pump its compression stroke; 
the return stroke is by the action of the spring. A cup leather 
on the piston acts as a suction valve on the return stroke. The 
Mercedes pump (Fig. 224), which is driven from the end of the 
camshaft, takes in air through ports uncovered by the piston near 
the end of the suction stroke. A relief valve is incorporated in 
the pump. 

Fuel pumps are made in many forms and are driven either 
from the engine or by small windmills. Sliding vane and gear 



FUEL SYSTEMS 



293 




Fig. 223. — Hispano-Suiza air pump. 



To Tank < 



Air Inlet Ports 



Relief Pressure 
Valve 




Fig. 224. — Mercedes air pump. 




Fig. 225. — Maybach fuel pump. 



294 



THE AIRPLANE ENGINE 



pumps (see p. 338) are often used and differ from the oil pumps 
only in smaller capacity. The compact duplex reciprocating 
pump of the Maybach engine (Fig. 225) is driven from a crank 
on the end of the oil-pump shaft through a yoke with a sliding 
bushing. Any leakage of gasoline past the plungers is into the 
crank chamber, which is filled with lubricating oil under pressure. 
Another method of avoiding the use of glands past which fuel 
leakage might occur is the employment of castor oil as the dis- 
placing medium. In the Benz engine (Fig. 227) the fuel pump 
is driven by worm gearing from the end of the inlet camshaft. 



,.-- Pressure Relief Valve 



To Carburetors 



Top of Gasoline ■■' 
Tank 




Tachometer Drive 



Sasoline Delivery to 

Pressure Reservoir 

in Main Tank and Hand 

Pump and Auxiliary 

Tanks 

— Outlet Check Yalve 

'"-Inlet Check Ya/ve 



'■ Gasoline Supply 
from Main and 
Auxilliary Tanks 



Fig. 226. — Benz pressure reservoir. 



Fig. 227. — Benz fuel pump. 



The lower portion of the cylinder is near the bottom of a chamber 
containing castor oil and the reciprocation of the piston produces 
a rise and fall of the castor oil in the annular space around the 
cylinder. The castor oil acts like an annular piston sucking in 
gasoline as its level falls and discharging it as the level rises. 
As the speed of the pump is slow (worm-gear reduction 10.75 to 
1) it is necessary to keep the discharged gasoline under air pres- 
sure during the suction stroke and this is accomplished by the 
use of a pressure reservoir (Fig. 226) located in the main fuel 
tank; the pressure reservoir also serves to damp out pressure 
pulsations. 



CHAPTER XII 

IGNITION 

Ignition is produced by the passage of an electric arc through 
the explosive mixture, at a time which varies somewhat with 
operating conditions, but in airplane practice is about 25 to 30 
deg. before dead center on the compression stroke. At the 
operating speed used in airplane engines the moving electrode 
of the make-and-break system is impracticable. The spark 
passes between stationary electrodes and is incorporated in 
" spark plugs" which are screwed into the cylinder head. For 
the production of the electric arc the following pieces of apparatus 
are necessary. 

1. A source of electric energy; this may be a primary or 
secondary (storage) battery, or more usually, a magneto. 

2. As the potential required to cause arcing is very large the 
low potential current generated in a battery or low-tension 
magneto has to be transformed into a high-potential current by 
the use of an induction coil; this is usually incorporated in the 
magneto. 

3. The current from a single source has to be sent in succession 
to each of several cylinders; this is accomplished by the use of a 
distributor which is located in the high-tension circuit. 

4. The distributor connects up the circuit to that cylinder in 
which ignition is next to occur and maintains that connection 
throughout a short period. The actual timing of the ignition 
within that period is controlled by a timer, breaker, or interrupter 
located in the low-tension circuit. 

Other minor but essential elements will be discussed later. 

Electric ignition systems utilize electro-magnetic phenomena. 
An electric current is induced whenever a conductor is moved 
through a magnetic field or when the magnetic field around a 
conductor is varied. The intensity of the induced current is 
proportional to the rate at which the conductor cuts the lines of 
magnetic force and to the number of coils cutting the lines of 
force. 

295 




296 THE AIRPLANE ENGINE 

The simplest kind of electric ignition system is shown in Fig. 
228; B is a source of current, N' a coil of wire surrounding an iron 
core (forming an electric magnet), S is a switch, timer, or other 
device for breaking the circuit at any desired moment. The 
magnetic flux is represented by the arrowed lines. If the switch, 
S, is opened the current falls to zero and N' is surrounded by a 
diminishing magnetic field; if S is closed 
N' is surrounded by a rising field. In both 
cases self-induction occurs in the coil N f 
and a current is generated in it, whose mag- 
nitude depends on the rate at which the 
magnetic field through N' changes and on 
the number of turns in the coil. The 
Fig. 228.— Inductance phenomena on closing and on opening S 

or spark coil ignition ., j.™ j. r\ i • j.i~ 

system. are quite different. On closing the cir- 

cuit, current can flow only after the switch 
is actually closed and the flow is opposed by the resistance of 
the circuit (which is small) and the self-induction pressure. 
When, Ihowever, S is opened, an air gap of great resistance is in- 
troduced into the circuit with the result that the current 
diminishes very rapidly and therefore establishes a high 
electromotive force by self-induction in N f . This electromotive 
force is sufficient to overcome the resistance of the small air gap 
formed at the instant of breaking contact and an arc is estab- 
lished across the gap. The resistance of the arc is considerably 
less than that of the air gap so that the current may continue 
to flow for a short time across a considerable arc. The more 
rapid the opening of the gap, the longer will be the arc. The 
energy for the arc is almost entirely the magnetic flux through 
the coil N f and is of comparatively small magnitude. 
If an additional or secondary coil N" be 



wound concentric with the primary coil N , 

as in Fig. 229, the same magnetic changes ^ B c ' 

will occur in both coils and an electro- T-E: 






motive force will be generated in N" FlG 2 29— High-ten- 
which is proportional to the number of sion or jump-spark igni- 
turns in the coil. When the number of lon sys em ' 
turns is very large, a high tension, sufficient to jump an air gap 
such as ab } will be produced. A coil wound as in Fig. 228 is called 
an inductance or spark coil. A coil with a double winding as 
in Fig. 229 is called an induction coil or jump-spark coil. 



IGNITION 



297 



The formation of an arc results in the vaporization or burn- 
ing of the metal of one of the points between which the arc 
springs and results in deterioration of that point. To reduce the 
arcing at S, sl condenser is shunted around it. The condenser 
consists of two conductors separated by insulating material; 
it is usually made of a large number of sheets of very thin metal, 
such as tin foil, separated by thin paraffined paper sheets. Every 
other sheet of metal extends to one side and the balance to the 
other. All the sheets of one side are connected to one terminal 
and the remainder to another. By connecting the condenser 
across the switch, S, (Fig. 229) the energy which would otherwise 
go into the formation of an arc is absorbed in the system. 



6 



«■ 



r i 




Fig. 230. — Circuit diagram of battery ignition system. 



The schematic arrangement of a battery ignition system for a 
four-cylinder engine is shown in Fig. 230. The primary circuit 
includes the battery, B, switch, S, primary winding on the induc- 
tion coil, I, the interrupter, breaker, or timer, T, which breaks the 
primary circuit whenever ignition is required, and the condenser, 
C, shunted around the timer to prevent arcing. The secondary 
circuit consists of the secondary winding of the induction coil, 
the distributor, D, and the spark plugs, p,p,p,p,; a safety spark 
gap, G, (see p. 310) is shunted on this circuit. The revolving arm 
of the distributor, D, establishes contacts successively with the 
four spark plugs in any desired order; the interrupter, T, breaks 
the primary circuit and the current thereby generated in the 
secondary circuit arcs across the spark plugs. The circuits are 
grounded as indicated. 

The Magneto. — Most airplane engines at the present day have 
magnetos as sources of electric current. A magneto differs 



298 



THE AIRPLANE ENGINE 



from a dynamo or electric generator in having permanent mag- 
nets in place of electro-magnets for the fields. 

In Fig. 231 is shown the action of an armature type magneto, 
consisting of pole pieces, N, S, which are permanent magnets, 
and an armature, A B, consisting of core and end pieces, revolving 
between the shoes of the pole pieces. The clearance ("air gap") 
between armature end pieces and magnet shoes is only about 
0.005 in. A coil is wound on the armature core, one end of the 
coil being grounded; the other end is carried away, insulated, 
through a collector ring and brush. As the armature revolves 
(being driven from the engine shaft) the lines of magnetic force 
take the successive directions indicated by the long arrows. The 
magnetic circuit is NABS for positions I and II. In the vertical 
position flux through the core ceases, and no current is generated 



/T\/T\ 




n m is: v 

Fig. 231. — Armature type magneto. 



in the coil. As the armature passes the vertical position, the 
circuit reverses to NBAS. This continues for 180 deg. more, 
when the original direction of flow is restored. The strength of 
the magnetic field influencing the armature coil is greatest at 
horizontal positions of the armature; but the rate of change of 
field strength is greater near the vertical positions, where the 
direction of magnetic flux is reversing itself. The air gap (in a 
construction like Fig. 231) is then large, so that the maximum 
effective rate of change occurs shortly after leaving positions II 
and V. Hence at these positions, twice in every revolution of 
the armature, the induced current reaches a maximum value, 
and is capable of producing a vigorous spark. 

Figure 232 shows the method of variation of the induced 
current with magneto position. Starting at position II, Fig. 231, 
the magnetic flux begins to diminish and has completely reversed 
itself by the time position III is reached. The duration of 
this period depends on the width of the armature end pieces. 
From position III to V there is practically no induced current. 



IGNITION 



299 



A high-tension magneto differs from that just described in 
that it has both primary and secondary coils wound on the 
same armature. Both coils link with the same magnetic circuit 
and therefore the armature becomes an induction coil and 
replaces the separate induction coil 
which would otherwise be necessary. 
Figure 233 shows a high-tension mag- 
neto. The two windings are shown 
with one terminal grounded to the 
machine. The primary coil is short- 
circuited by the contact at the inter- 
rupter, at M, until the proper moment, 
when it is opened suddenly and the 
induced high-tension current goes 
through the distributor to one of the 
spark plugs. The magneto and inter- 
rupter must be properly synchronized 
so that the break occurs when the primary e.m.f. is a maximum. 
The switch, when closed, short-circuits the primary circuit and 
thereby prevents the building up of a high-tension current in the 
secondary circuit, and so shuts off the ignition. 

Of the elements shown in Fig. 233 the condenser and inter- 
rupter are usually incorporated in the actual construction of the 




it 



iff 



Position of Armature 

Fig. 232. — Induced current in 
armature type magneto. 



Spark P/ujs 
r 




Condenser . 
Ground" 
Fig. 233. — Circuit diagram of high-tension magneto ignition system. 



magneto. The distributor may also be incorporated when the 
magneto speed is one-half the engine speed. 

The ordinary construction of a magneto with revolving armature 
gives sparks at 180-deg. intervals corresponding to the positions 



300 



THE AIRPLANE ENGINE 



(II and V, Fig. 231) of maximum induced current. With Vee 
type engines it may be necessary to have unequal time in- 
tervals between sparks; for example, with a two-cylinder 45-deg. 
Vee engine the sparks instead of occurring at 180-deg. rota- 
tion of the armature should occur alternately at 157H"deg. 
and 202 M-deg. intervals. This unequal interval can be obtained 
in various ways. In one of the constructions of the Bosch 
Magneto Co. the armature end-piece is cut away on opposite 




^5 

I ir m 

Fig. 234. — Magneto with unequal firing intervals (Bosch) . 

sides of each half of the core so as to increase the air gap and 
the tips of the pole shoes are also cut away on diagonally opposite 
halves of the two poles so as to make the positions of maximum 
induced current (II, Fig. 231) come earlier. The construction 
and operation are illustrated in Fig. 234. The large air gap, B, 
effectively cuts off the lines of force. Maximum induced current 
will occur shortly after the armature has left the trailing pole 
tips C - D (position II) and also after the armature has left 
the trailing pole tips E-F (position IV). These two positions 




5 N 



m 



S N 



w 



S N 



Fig. 235. — Inductor magneto. 

are made less than 180 deg. apart as a result of cutting away tips 
of the pole pieces at E and F. 

The magneto with revolving armature has to be provided 
with insulated moving wires, collector rings, brushes, and moving 
contacts to convey the induced current from the armature to 
the stationary conductors. To avoid this complication a rotor 
or inductor type of magneto, with stationary windings, is often 
used. Figure 235 shows a construction with a rotating element, 



IGNITION 



301 



or inductor, consisting of two cylindrical segments of soft iron; 
all the rest of the magneto is stationary. The magnetic condition 
of the armature core depends on the position of the inductor. 
In the positions A and C the segments form a magnetic bridge 
between the magnet poles and the heads of the armature core; in 






Fig. 236. 



90° 180° 

Rotation of Inductor «— ^ 
Induced current in inductor magneto. 




these positions the magnetic flux is a maximum. In passing the 
positions B and D the magnetic lines are abruptly changed in 
direction and a vigorous induced current is set up. The reversal 
takes place four times per revolution of the inductor and succeed- 
ing reversals give current in opposite directions. This inductor 
magneto can give twice as many ignitions 
per revolution and consequently has to be 
rotated only half as fast as the rotating 
armature type of magneto. All the elec- 
trical connections are stationary. Typical 
current curves are shown in Fig. 236. 

Another construction of inductor magneto 
is shown in Fig. 237. The rotor is a steel 
shaft carrying two laminated soft-iron arms 
with a space between them which is occupied 
by the stationary winding (not shown) . The 
arms project on opposite sides of the shaft 
and are of such radius as to give the smallest 
practicable air gap between them and the 
pole shoes. The magnetic flux in the 
position shown is from N to R, then back 
along the shaft to the other arm and so 
to S. When the inductor has rotated 180 
deg. from the position shown, the flux will be from N to the rear 
arm, then forward along the shaft to R and S. The flux through 
the shaft is reversed twice every revolution and induces a cur- 
rent in the winding around the middle length of the shaft. 



1 



^ 



i 



i 



Fig. 237. — Inductor 
magneto with two 
arms. 



302 



THE AIRPLANE ENGINE 



The Dixie magneto uses a different type of inductor. The 
rotor, Fig. 238, consists of two revolving wings, N and S, sepa- 
rated by a bronze center-piece, B. Each wing is always in contact 
with one pole of the magneto (Fig. 239) and consequently keeps 
the polarity of that pole. The rotor is surrounded by the field 
structure, shown in Fig. 240, which carries laminated pole exten- 
sions on which the winding with its core is mounted. As the 



a 



3 



o B o 



F^ 



,. lyrji 



Fig. 238. — Rotating element of Dixie magneto. 

rotor revolves the direction of magnetic flux through the core 
changes twice every revolution. 

Construction of Magnetos. — The constructive features of a 
Bosch high-tension magneto of the rotating armature type are 
shown in Fig. 241. The armature rotates at engine speed and 
gives two electrical impulses per revolution. The distributor 
is geared to the contact breaker and rotates at half its speed. 

The end of the primary winding is connected to the brass 




G33: 



Fig. 239. — Diagrammatic outline 
of Dixie magneto. 




Fig. 240. — Field structure 
of Dixie magneto. 



plate, 1. In the center of this plate is screwed the fastening 
screw, 2, which serves, in the first place, to hold the contact 
breaker in its position, and, in the second, to conduct the primary 
current to the platinum screw block, 3, of the contact breaker. 
Screw, 2, and screw block, 3, are insulated from the contact 
breaker disc, 4, which has metallic connection with the armature 
core. The platinum screw, 5, goes through the screw block, 3. 
Pressed against this platinum screw by means of a spring, 6, is the 



IGNITION 



303 



contact-breaker lever, 7, which is connected to the armature core 
and to the beginning of the primary winding. The primary 
winding is therefore short-circuited as long as lever, 7, is in 
contact with the platinum screw, 5. The circuit is interrupted 





Lonqi+udinal Sec+i'on 



Rear View 



Fig. 241. — Constructive features of Bosch high-tension magneto. 

when the lever is rocked. A condenser, 8, is connected in parallel 
with the gap thus formed. 

The end of the secondary winding leads to the slip ring, 9, 
on which slides a carbon brush, 10, which is insulated from the 




Sparking P/uqs 



■ Primary Winding 
Secondary Winding 

■ Frame 



) - 5afef S 5 P ar ^ a PDisi-ribu-h i 





W) 



Con+acf- Breaker Disk 
Fig. 242. — Wiring diagram for Bosch high-tension magneto ignition system. 

magneto frame by means of the carbon holder, 11. From the 
brush, 10, the secondary current is conducted to the connecting 
bridge, 12, fitted with a contact-carbon brush, 13, and through 
the rotating distributor piece, 14, which carries a distributor 
carbon, 15, to the distributor disk, 16. 



304 



THE AIRPLANE ENGINE 



In the distributor disc, 16, are embedded four metal segments, 
17. During the rotation of the distributor carbon, 15, the latter 
makes contact with the respective segments, and connects the 
secondary current with one of the contacts. 

The contact breaker is fitted into the rear end of the armature 
spindle, which is bored out and provided with a keyway. The 
short-circuiting and interrupting of the primary circuit is effected 
by means of the contact-breaker lever, 7, and the fiber rollers, 19. 
As long as the lever, 7, is pressed against the contact screw, 5, the pri- 
mary circuit is short-circuited, and the rocking of the levers by the 
fiber rollers, 19, breaks the primary circuit; at the same moment 
ignition takes place. The distance between the platinum points, 
when the lever is lifted on the fiber rollers, must not exceed 0.5 
mm. (approximately }io in.). This distance may be adjusted by 
means of the screw, 5. 





Fig. 243. — Bosch interrupter-distributor. 

Another type of Bosch interrupter-distributor is shown in 
Fig. 243. This is connected at 1 to the engine cam shaft and 
carries a cam, 15, which has as many lobes as there are cylinders 
to the engine; four lobes are shown. The interrupter lever, 16, 
is held against this cam by the spring, 17, and when the rubbiDg 
plate of the lever is between two lobes the platinum points, 18 
and 19, are in contact; the contact is interrupted at the passage 
of each lobe and the primary circuit is thereby broken. The 
distributor rotor is on the same shaft and carries a rectangular 
brass tube in which is located the carbon brush, 11. This brush 
sweeps the cylinder cavity in the distributor body, 7, and the con- 
tacts, 8, which are as numerous as the cylinders. The central 
carbon brush, 10, keeps contact with the rectangular brass tube. 
Adjustment of the interrupter is by rotation of the whole dis- 
tributor through the timing arm, 6. 



IGNITION 305 

Permanent magnets are of steel alloyed with 5 per cent of 
tungsten or with chromium. All parts at which sparks may 
occur (breaker, distributor, safety gap) should be enclosed to 
reduce the fire risk. 

The distributor speed is hah the engine speed on all stationary 
four-cycle engines; the distributor may either be incorporated in 
the magneto or be driven direct by the cam shaft. The Dixie 
distributor for an 8-cylinder engine is shown in Fig. 244. The 
rotor carries two carbon brushes, which in an 8-cylinder engine 
are 180 deg. — 45 deg. = 135 deg. apart, and in a 12-cylinder engine 
are 180 deg. — 30 deg. = 150 deg. apart. These brushes are not 
in the same plane of revolution. In the plane of the outer brush 



Carbon 
Brushes 




Co/lector Brushes 
Fig. 244. — Dixie distributor for 8-cylinder engine. 

are located four metal segments embedded in the insulating 
distributor block; the other four segments are located immedi- 
ately behind the first four in the plane of the second carbon 
brush. Contacts are made each 45 deg. of rotation of the 
distributor rotor. The collector brushes are in continuous 
contact with the secondary circuit and with the carbon brushes. 
When rubbing contact is employed in a distributor a deposit 
of carbon from the brush will be left on the distributor block 
which must be cleared off periodically. To avoid this a gap 
distributor is sometimes used with a nickel point and a small 
air gap (from 0.01 to 0.02 in.) across which the current arcs. 
The use of a gap distributor has the additional advantage of 
increasing the secondary voltage when the spark plug has its 
resistance lowered either by carbon deposit or high temperature, 

20 



306 



THE AIRPLANE ENGINE 



and thereby giving a spark under conditions in which it would 
otherwise fail (see p. 310). 

The spark advance in airplane engines is generally fixed at 
about 30 deg. A slightly greater advance is desirable at high 
altitudes, but the advantage from its use is so slight and the 
complication of an additional control so undesirable that spark 
control is not used in airplane practice. Spark adjustment is 
obtained by adjusting the breaker. A corresponding adjustment 
of the magneto is sometimes provided. Figure 245 shows the 
adjustable driving gear arrangement of the magneto of the King- 
Bugatti engine. The bevel gear on the magneto shaft is fitted 




Packed with Soff~6rease 
Fig. 245. — Adjustable driving gear for Bugatti engine magneto. 



on a tap er with a key. The gear has eight keyways, so spaced 
that the magneto timing may be set within 1J^ deg. of any 
de sired position. The gear which adjusts the spark advance has 
four internal spiral grooves sliding over splines on the sleeve, 
which is keyed to the driving shaft, but may be moved along the 
shaft by a lever. Movement of this sleeve revolves the magneto 
driving gear with relation to the shaft-driving gear. 

Adaptation to Engine, — In four-cycle engines having n cylinders, 

Tt 

there are - sparks necessary per engine revolution. If these are 
supplied by a magneto giving m sparks per magneto revolution, 



the ratio of magneto speed to engine speed is 



2m' 



Since m 



v aries from one to four (the interrupter may work only once per 



IGNITION 307 

revolution even though conditions are right twice), the speed 

fit 71 

ratio varies from - to -. The great majority of magnetos give 

2 o 

71 

two sparks per revolution, and run at -i times engine speed. A 

multiple-spark magneto runs at relatively low speed. 

In an engine with equal ignition intervals, one magneto may 
serve any number of cylinders. Thus a two-spark magneto for 
15 cylinders would run at 3% times engine speed: the same 
magneto for 9 cylinders runs at 2J4 times engine speed. With 
unequal ignition intervals, a special magneto (see page 300) or a 
plurality of magnetos must be employed. 

The cycle of operations of a jump-spark ignition system 1 
can be considered as consisting of five periods. For quan- 
titative values a typical magneto may be considered to have the 
constants given in the following table. 

Constants of Typical Magneto 

Primary turns (JVi) 160 

Secondary turns (2V 2 ) 8 , 000 

Ratio of turns (n) 50 : 1 

Primary resistance (Ri) 0.5 ohm 

Secondary resistance (R 2 ) 2,500 ohms 

Primary inductance (Li) 0. 015 henry 

Mutual inductance (M) . 74 henry 

Secondary inductance (L 2 ) 36 henrys 

Primary condenser (Ci) 0.2 microfarad 

Secondary (distributed) capacity (C 2 ) 50 micro-microfarads 

Normal speed of operation 2,000 r.p.m. 

Primary current at break (7 b ) 4 amperes 

Maximum current in spark . 075 amperes 

Breakdown voltage of gap . 5 , 000 volts 

Sustaining voltage of gap 600 volts 

Period I includes the building up of current in the primary 
winding as a result of either the impressed voltage from a battery 
or the voltage generated by the rotation of a magneto armature. 
During this period the breaker or interrupter is closed and the 
armature rotates from the position of maximum flux to the 
position where the interrupter opens. For the typical magneto 
this corresponds to about 100-deg. rotation and lasts 0.008 sec. 
at 2,000 r.p.m.; the current builds up to 4 amperes. Typical 

1 See Report 58, 5th Annual Report, Nat. Adv. Comm. Aeronautics. 



308 



THE AIRPLANE ENGINE 



curves for armature flux are shown in Fig. 246, in which A shows 
the flux with open primary circuit and is due to the permanent 
magnetos only. B and C give the total flux under normal operat- 
ing conditions at 500 and 2,000 r.p.m. respectively. 

Period II is the very short period (about 0.00002 sec.) extend- 
ing from the opening of the interrupter to the breakdown of the 
spark gap in the engine. During this period, the magnetic 
energy of the coil is in part transferred into electrostatic energy 
and charges the condenser and the capacity of the secondary 
circuit. The primary current flowing into the condenser against 



40 

30 

20 

10 



-10 

-20 

-30 

-40 























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S£ 


Y 










/ 


if 


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Y 


■eak 








A 


8 C 




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i 




















































i 


r 


\ 




1 




/ 


Break 






\ 


\ 




I 


I 


7 








N 


w. 


' 


c 


Break 





















50 50 100 150 200 250 300 350 400 
Angle of Rotation (Degrees) 

Fig. 246. — Typical curves for armature flux of magneto. 



a constantly increasing e.m.f. will decrease at a constantly 
increasing rate. The decrease in magnetic flux resulting from 
this decrease of primary current generates an e.m.f. in the sec- 
ondary windings which in turn sends a charging current into the 
distributed capacity of the secondary circuit. If the spark gap 
were not present this process would continue until the magnetic 
energy had been entirely converted into electrostatic energy; 
the maximum voltage which would be reached in the typical 
magneto would be about 70,000 volts. As a result of loss of 
energy due to resistances, eddy currents in the iron core, etc., 



IGNITION 



309 



this maximum voltage is greatly reduced. The curves of Fig. 
247 show the rate of rise of secondary voltage as calculated; 
(A) with no energy loss except that in the resistance of the wind- 
ings, (B) with the usual eddy currents. It is assumed that the 
spark gap does not break down and there is no arcing at the 
interrupter points. 

Period III is the very short period (about 0.00005 sec.) begin- 
ning at the instant at which the spark gap breaks down and 
lasting until a steady arc is established. When this gap has 
broken down it affords a conducting path into which the charged 
secondary capacity discharges. As the secondary is now short- 
circuited by the arc the current increases rapidly. The energy 



40,000 



£ 30,000 



[p 20,000 
n 

§ 10,000 

















A % 
























































































£- 











































































































































10 40 60 80 100 120 140 

Time After Break"(Millionthsof a Second) 

Fig. 247. — Typical curves for secondary voltage of magneto/] 



discharged into the gap during this time is about 0.002 joule, 
which is just about sufficient to ignite the explosive mixture 
(see p. 312). 

Period IV extends from the establishment of the gap to the 
extinction of the spark. During this time there is a steady 
discharge which lasts for a considerable time (0.003 sec. for a 
5-mra. spark gap in air). The cessation of the arc is due usually 
to the exhaustion of the energy supply, although occasionally 
it may be extinguished by the closing of the interrupter if the 
r.p.m. is very high or the spark gap very short. 

Period V covers the remainder of the cycle during which 
time both circuits are practically free from current. 

Of these periods, II is the most important, as it determines 
whether or not a spark passes at all; the distributed capacity 
of the secondary circuit is of great moment in determining the 
maximum voltage in this period. While 5,000 to 6,000 volts 



310 



THE AIRPLANE ENGINE 



is usually required to jump the gap, it may be much increased 
by oil films on the points and a cold cylinder. If the spark plug 
is fouled with a conducting film of carbon some of the energy 
will be drained by leakage during Period II. 

A typical oscillograph showing the variations of the primary 
and secondary circuits is given in Fig. 248. The maximum 
current delivered to the secondary circuit is usually from 0.05 to 
0.10 amperes. 

A safety spark gap is sometimes shunted on the secondary 
circuit (Figs. 230, 241, and 242) to prevent the formation of 

excessive voltages, and the 
consequent possible break- 
ing down of the insulation, 
in case the secondary cir- 
cuit is open. This would 
occur when a spark plug is 
being tested out of the cyl- 
inder and is not grounded. 
The width of the safety 
gap is from %q to % in., 
the higher value being used 
for high compression in the 
engine. 

The air surrounding the safety spark gap becomes ionized 
and ozone is liberated. If the air is confined the ozone will 
rust adjacent steel parts and slowly decompose organic insulating 
materials. A rotary safety gap with one electrode on the gear 
driving the distributor and the other integral with the distribut- 
ing metal electrode is used sometimes; the air is churned up and 
expelled through a suitable gauze window. 

A series or subsidiary spark gap is frequently used in order 
to maintain sparking even when the spark plugs are fouled. 
The series gap is placed in the connection between the plug 
and the magneto. 1 Investigations at the Bureau of Standards 
show that it is possible, by the use of a series gap, on an average 
ignition system, to spark a plug having a resistance lowered to 
only 4,000 ohms by fouling. At least 100,000 ohms insulation 
resistance is ordinarily necessary at the plug if a series gap is not 
used. For example, with secondary current limited to 0.08 

1 For elementary theory and results of tests see Report 57, 5th Annual 
Report, Nat. Adv. Comm. Aeronautics. 




Fig. 248. — Typical oscillograph of primary 
and secondary currents in magneto. 



IGNITION 



311 



amperes and insulation resistance of 50,000 ohms the maximum 
voltage across the air gap is 0.08 X 50,000 = 4,000 volts. This 
is not sufficient; 6,000 volts is usually required. 

The efficacy of the series spark gap is well shown in Fig. 249, 
giving the results of some tests by Young and Warren. 1 The 
four resistances indicated were put in parallel with the spark plug 
to simulate different degrees of fouling. With each resistance the 
length of the main gap was varied while the series gap was 
kept constant. The curves show the maximum length of the 
main gap at which sparking oc- 
curred. With a parallel resistance 
of 58,000 ohms the main gap could 
be increased from 0.9 mm. to 3.4 
mm. as the series gap was increased 
from zero to 0.02 in. The voltage 
across the main gap was also 
measured and was found to in- 
crease from 2,200 to 3,600 by the 
introduction of a 0.02-in. series gap, 
when the shunted resistance was 
112,000 ohms. 

The series gap is sometimes in- 
tegral with the plug; sometimes 
at the plug but not integral with 
it; sometimes in the distributor. 
The amount of the gap should be 
of fouling of the plug. For 







~F 


--250, 000 












f?= 


■ 140, 000 








, -** 








/ 


s 




^Wg 














r R 


'=5d 


POL 


7 

























■ E 6 

D 
ID 

|4 



0.02 0.04 0.06 0.08 
—^ Length of Auxiliary Spark 
Gap in Inches 

Fig. 249. — Effect of auxiliary spark 
gap on length of main gap. 



variable to suit the degree 
maximum effectiveness the gap 
should be at, or integral with, the spark plug. These desider- 
ata are conflicting. It is not practicable to adjust subsidiary 
spark gaps at each spark plug, while the engine is operating; 
it is easily possible to have a single adjustable spark gap at the 
distributor, but it cannot be adjusted to suit the different degrees 
of fouling in the different cylinders which it serves. 

The effect of temperature and pressure on sparking voltage 
has been investigated at the Bureau of Standards. 2 Sparking 
voltage is a linear function of the density of the gas and depends 
on pressure and temperature only as they affect the density. 
For a typical spark plug with 0.5 mm. gap the sparking voltage 

ia The Process of Ignition," The Automobile Engineer, March, 1920. 
2 See Report 54, 5th Annual Report, National Advisory Committee on Aero- 
nautics. 



312 



THE AIRPLANE ENGINE 



in air varies from 2,800 volts at atmospheric density to 9,400 
volts at a density five times as great. Other measurements 
indicate that the sparking voltage in an explosive mixture of 
gasoline in air is about 10 per cent less than in pure air and 
that the change in voltage is proportional to the percentage of 
gasoline present. Figure 250 shows the observed sparking 
voltage for plugs with different gaps; No. 1, 1.8 mm. (0.071 
in.); No. 2, 1.2 mm. (0.047 in.); No. 3, 2.2 mm, (0.086 in.); 
No. 4, 0.5 mm. (0.020 in.). The voltage required for a spark 
plug set at 0.5 mm., in an aviation engine of moderate compres- 
sion, is about 6,000 volts. 



en m 

2 ]i 



£ 10 

CO 



















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Y/ 
























6 


V 
























































































































A 










































































































































Density Referred to Ai r at 1 Atm , and 273 Decj. Abs.,Cent. 
Fig. 250. — Sparking voltages for plugs with different gaps at various air densities. 



The sparking voltage is not affected appreciably by the 
material of the electrodes but is diminished by the use of finer 
points. The dimensions of the points are, however, determined 
by considerations of mechanical strength and durability. 

The " fatness" of a spark has no influence at all on the power 
developed in an engine. If the current is sufficient to charge 
the plug and its connections to the sparking potential, the 
maximum engine power will be developed. The energy repre- 
sented by that condition is usually about 0.002 joule; the energy 
per spark varies from 0.03 joule in battery systems up to 0.16 
joule in the more powerful magnetos. The excess of energy 
above that necessary for ignition has no discernible effect on the 
power developed. 

Battery Ignition. — In a storage cell, electrical energy is stored 
as chemical energy but returns to electrical energy when the cell 



IGNITION 313 

is connected to supply an external circuit. The desired voltage 
is obtained by connecting a sufficient number of cells in series, 
forming a storage battery. 

The chemical reaction in a lead cell may be expressed by the 
following equation: 

Charge 

< 

Pb0 2 + Pb + 2H 2 SOt = 2PbS0 4 + 2H 2 

Positive and 
Positive plate Negative plate Electrolyte negative plates Electrolyte 

Discharge 

Discharge results in the formation of lead sulphate; charging 
restores the plates to their original conditions of lead sponge 
(negative) and lead peroxide (positive). 

The voltage of a fully charged idle cell is 2.05 to 2.10 volts. 
Discharge lowers the voltage in proportion to the current flowing. 
Complete discharge is reached at 1.7 volts, at the normal dis- 
charge rate fixed by the manufacturer. The capacity of a 
battery is expressed in ampere-hours at normal discharge rate; 
the capacity increases as the discharge rate decreases. The 
maximum discharge rate falls as the temperature decreases. 

The acid or electrolyte is an aqueous solution of density 1.255 
resulting from the addition of 1 part of sulphuric acid of sp. gr. 
1.84 to 4J^j parts (by volume) of distilled water. The strength 
and density of the solution fall as discharge progresses; when 
the density falls below 1.2 the cell needs recharging. 

Self-sustaining battery systems require a generator for 
recharging. The system may then be regarded as a generator 
system on which the battery "floats." The generator furnishes 
low-tension direct current, and must have a commutator and 
brushes. In starting, or at low speed (up to 650 r.p.m.), ignition 
current comes from the battery. At some definite speed, say 
650 r.p.m. of the engine, the generator begins to supply the 
ignition current. Its rate of delivery is then considerably in 
excess of that needed for ignition and the surplus goes to the 
battery. The recharging rate has a maximum value of 10 
amperes when the battery is nearly discharged. The delivery 
voltage of the generator may be automatically controlled by a 
potential regulator. 

The outlines of the Liberty engine battery ignition system 



314 



THE AIRPLANE ENGINE 



are shown in Fig. 251. Figure 252 shows the circuit diagram. 
It includes a low-voltage generator in connection with a storage 



Left Distributor Generator 

/ Gen. Arm. Gen. Field 




Gen Field Gen. Arm. ' Bait. 

Fig. 251. — Liberty engine ignition system. 

battery of light weight and small liquid content. Through a 
switch, the current is sent to two distributors which are mounted 




Non-inductive 
Resistance 



Reverse Coif 



Generator Shunt 
Field Coi! 



/ 



jr-Vo/fcrcfe Coil 



Voltage Regulator 



Resistance 
Unit-- 




Switch 



Storage 

-Resistance Baffer J/ " 

Unit. 
■Blades Separately 
Insulated 




Left Distributor Right Distributor 

Fig. 252. — Liberty engine circuit diagram. 

on the camshaft housings and are direct-driven by the camshafts. 
Right-hand distributors supply distributor-end plugs and left- 



IGNITION 



315 



Brush- 



Armature 



FizldCoils 



Lower End 
Hous/ncf. 



Generator Armature 
• Terminal 



'Commutator 



Upper End 
Housing 



hand distributors supply propeller-end plugs, there being two 
plugs to each cylinder. The entire system, exclusive of plugs 
and wiring, weighs 35 lb., and consists of generator, battery, 
switch, voltage regulator, and two distributors. 

The generator is shown in Fig. 253. It is four-pole, shunt- 
wound, 43^ in. diameter, S}^ m - high, and weighs ll^i lb. It 
is mounted with its shaft 
vertical on top of the 
crankcase between the 
two rows of cylinders at 
the rear end. The arma- 
ture shaft extends down- 
ward into the crankcase, 
where it is driven by the 
auxiliary gearing which 
also drives the vertical 
shaft. Its speed is 1.5 
crankshaft speed. The 
end housings of the gen- 
erator field frame are of 
cast aluminum. Gearing 
from the armature shaft 
forms the tachometer 
drive. The upper hous- 
ing contains the ball 
bearing for the shaft, 
which is the only bearing 
in the generator proper. 
This housing also sup- 
ports the four brush 
holders: two positive brushes 
grounded to the frame. 

The ground side of the field is grounded through the voltage 
regulator, (Fig. 252), the generator voltage being determined by 
the amount of current flowing through the field, which in turn is 
controlled by the regulator. The armature has 21 slots and is 
wave-wound. Insulation between commutator segments is slotted 
down J<32 in. below the surface of the copper bars. The shaft is 
hollow, and ground through the bearing. The maximum generator 
voltage is 10 to 10^2 volts. A current of 5 to 6 amperes may be 
carried without overheating. 




Sphhed ' fnd 
of Armature 
Shaft 



Fig. 253. — Liberty engine generator. 



insulated, and two negative, 



316 THE AIRPLANE ENGINE 

The voltage regulator keeps the voltage constant at all speeds 
above 650 r.p.m. of the engine. It weighs lJ-£ lb- and is mounted 
in a cast aluminum cup on the back of the dash behind the switch. 
It consists of a soft-iron core over which a pivoted iron armature 
is so mounted as to be normally held away from the core by an 
adjustable tension spring. When so held, the generator field 
current passes through a tungsten contact point on the armature 
to the ground. 

The core carries three windings (Fig. 252). The voltage 
winding is of fine wire leading from the positive terminal of the 
generator armature to the ground. The generator voltage is 
impressed on this winding. Increase in this voltage increases the 
core magnetism and opens the contact gap of the regulator 
armature. This cuts off the direct flow of field current and 
decreases the armature voltage of the generator. The reverse 
winding is superimposed on the voltage winding and is also of fine 
wire, wound in a reverse direction. The non-inductive winding 
consists of resistance wire wound so as to produce no magnetic 
effect on the core and so as to be itself free from induction due to 
changing core-flux. These two windings are connected in 
parallel from the regulator armature contact point to the ground : 
they form a permanent high-resistance ground for the field 
current when the contact is open. The reverse winding rap- 
idly destroys residual magnetism and enables the spring again 
to close the contacts, which in regular operation vibrate 
rapidly. 

The battery is designed for light weight and no leakage. It 
is 7 by 4 by 5% in. and weighs 10J4 lb. It can provide 3 amperes 
for 3 hr., which is sufficient energy for dual ignition on 12 cylin- 
ders. It floats on the line and normally supplies current only 
when the engine speed is under 650 r.p.m. The ammeter shows 
whether the battery is charging or discharging. Charging is 
automatic at speeds above 650 r.p.m. The hard-rubber battery 
jar has four compartments or cells, each cell (Fig. 254) con- 
taining 3+ and 4— plates, burned to connecting straps and 
separated by perforated rubber with wood. Plates are 3 by 3 
in., and rest on %-in. bottom ribs. Above the top of each plate 
is a flat sealing or baffle plate of hard rubber. The top of each 
cell is further sealed by a rubber cap through which the lead 
terminal posts extend. These also are sealed by gaskets or 
by burning. 



IGNITION 



317 



In normal position, the electrolyte completely fills the plate 
compartments. When the battery is turned upside down, the 
electrolyte seeps through small holes to the compartment which 
is normally above the plates. This compartment is of such capa- 
city as to hold all of the fluid in the cell, without danger of over- 
flow at the vent plug. 

The switch, Fig. 252, is located on the dash and weighs 1 lb. 
It is built on a Bakelite base. The circuits are controlled by 

Top Connector ,-■■ Vent Pluj Washer .- Posit/Ye Term/haf 
// Negative Term/haf 

Vin+PIucj j^ Positive and 

_ 'ative Piats 
Assembly 




Fig. 254. — Section of storage cell. 



two aluminum switch-levers operating spring-bronze contact 
fingers which connect with contacts molded in the base. The left 
lever supplies the left distributor, the right lever the right dis- 
tributor. The engine is started on one distributor, and both 
levers are switched on only when the speed is above 650 r.p.m. 
(To start on both distributors would require that the battery 
supply two sets of plugs and would also waste battery current 
through the generator.) Resistance coils are mounted on the 
back of the switch in series with the distributor circuits. These 
prevent an excessive flow of current should the switch be left on 
with the engine idle. The switch has four external connections: 



318 



THE AIRPLANE ENGINE 



positive battery, generator armature, and two to distributors. 
Two 12-cylinder distributors (Fig. 255) are used, each supply- 
ing one plug in all cylinders. Each weighs 5 L ^ lb. and is 7% in. 
diameter by 5% in. high. They are mounted one on each of 
the overhead camshafts. The transformer coils and breaker 
mechanism are incorporated. The Bakelite distributor head 
forms a cover for the breaker mechanism and a seal for the coil. 
The moving contact is through a soft carbon brush bearing on 
terminals molded in a hard-rubber track. 



Felt Washer. 



Driving 

F/arrae, 




Retainer 




Rotor Arm 



Rotor Brush 

Low Tension 
Lead 'to 5 'wit- h 
Fig. 255. — Liberty engine distributor. 




Fig. 256. — Liberty engine 
breaker mechanism. 



The breaker mechanism is operated by a 12-lobed cam, having 
lobes spaced 22^ deg. and 37M deg. (12-cylinder engine). 
Tungsten contact points are used. Two main-circuit breakers 
a and b, Fig. 256, connected in parallel, are provided and are 
timed to operate simultaneously; the duplication is a precaution- 
ary measure. The auxiliary -circuit breaker, c, is provided to 
prevent the production of a spark when the engine is " rocked" 
or turned backward. This auxiliary breaker is connected in 
parallel with the other two through a resistance unit (Fig. 252) 
which reduces the amount of current flowing through it and is so 
timed that it opens slightly before the other two when the engine 
is turned in a forward direction. The opening of the main 



IGNITION 319 

breakers then results in the production of a spark. When the 
engine is turned in a backward direction the two main breakers 
open first and no spark is produced due to the fact that the 
current continues to flow through the coil through the auxiliary 
breaker but in diminished quantity due to the resistance unit. 
By the time the circuit is opened at the auxiliary breaker the 
intensity of the magnetic field of the coil has weakened to such 
an extent that no spark is produced. The whole breaker mechan- 
ism may be revolved to advance or retard the spark. 

Spark Plugs. — The conditions to which the spark plug is 
subjected in aviation engines are difficult to meet. The require- 
ments are: 

1. The maintenance of a gap having a breakdown voltage 
of about 6,000 volts. 

2. The maintenance of an insulation resistance of at least 
100,000 ohms. 

3. Practically complete gas tightness. 

These conditions must be maintained under pressures of 500 to 
600 lb. per square inch while immersed in a medium which alter- 
nates, 15 cycles per second, in temperature between 50° and 2,500° 
C. and in an atmosphere which tends to deposit soot, and possibly 
oil, on the surface of the insulator. The inner end of the insulator 
and central electrode may have an average temperature of 900°C. 
while the body of the insulator, well up in the shell, is in contact 
with a jacket containing water at 70°C. For successful operation 
the insulating surface must remain clean, the insulator must not 
fracture or disintegrate under the varying temperatures and no 
part of the plug must become hot enough to cause preignitions 
of the charge. 

The shell is of steel with standard thread. The S. A. E. stand- 
ard plug dimensions are: outside diameter, 18.2 mm.; pitch 
diameter, 17.22 mm. +0.02 mm.; root diameter, 16.09 mm.; 16.9 
threads per inch; 1.5 mm. pitch. In order to keep down its 
temperature it is often made with fins for radiating heat. Investi- 
gations at the Bureau of Standards 1 have shown that brass shells 
average from 90° to 270°F. hotter than steel shells. 

The most common insulating materials are mica and porcelain. 
Other materials used are fused quartz, steatite and molded mater- 
ials such as Bakelite and Condensite. Porcelain of the highest 
grade is an excellent material except for its brittleness and conse- 

1 Report 52, 5th Annual Report, Nat. Adv. Comm. for Aeronautics. 



320 THE AIRPLANE ENGINE 

quent liability to fracture either from temperature or mechanical 
effects. Mica, while free from this trouble, is more likely to foul 
in consequence of its rough surface. Fused quartz is free from 
both the above objections. 

The insulation resistance of these materials diminishes with rise 
in temperature. At a temperature of 900°F. the order of merit 
of the insulating materials is mica, quartz, steatite, porcelain. 
Porcelain plugs which show a resistance of millions of ohms 
when cold may have a resistance of only 100,000 ohms at 900°F. 

The most common source of failure in spark plugs is fouling. 1 
It causes more than 50 per cent of spark plug troubles and is 
most serious at high altitudes. Fouling is due to the deposit of a 
layer of carbon and causes a short circuit. The carbon deposit 
results from either or both of two causes: (a) The chilling of 
the flame by a cool portion of the plug and the consequent 
incomplete combustion; this effect is particularly common when 
the mixture is overrich and is of frequent occurrence at high 
altitudes with an imperfectly compensated carburetor, (b) 
The decomposition of lubricating oil which is splashed on heated 
portions of the insulator. The oil itself is an insulator and 
when it wets a layer of soot in the plug it makes it an insulating 
layer. Such a deposit chars and becomes more and more con- 
ducting. The oil acts as a binding material and also increases 
the rate of deposition of soot since the carbon particles in the 
flame adhere to it readily. The conduction through the deposit 
seems to take place through a narrow path where the oil film 
between the particles has been broken down by electric stress. 
Such fouling causes misfiring. 

A method of attempting to reduce the deposit of carbon is to 
keep the insulation at such high temperature that all carbon 
deposit is burned off. This can be accomplished by making the 
insulator with petticoats, ridges or other projections, but there 
results the danger of preignition, particularly in high compression 
engines or in those that are not well cooled. Another method is 
to shield the insulator with a metal baffle plate which protects 
it from oil spray, but if this is done the flame does not get good 
access to the insulator and any deposit which has formed will 
have very little opportunity of being burned away. The use 
of a series gap (p. 310) is useful in maintaining firing after the 
plug is fouled. 

1 Report 52, 5th Annual Report, Nat. Adv. Comm. for Aeronautics. 



IGNITION 321 

Fouling with oil, either in the form of a surface film over the 
electrodes or as a drop between the points, will often prevent 
firing. The breakdown strength of oil is several times that of 
air, and the voltage required may easily exceed that which the 
ignition system is capable of delivering. The trouble is intensi- 
fied if the insulation of the plug is at the same time reduced by a 
layer of soot, thereby diminishing the maximum voltage which 
the system can develop. 

The oil trouble usually occurs on starting but may also be 
met when the plane is recovering from a long glide during which 
the engine is turning over slowly and pumping oil into the 
cylinders. It may sometimes be identified by the sparking of 
the safety gap (see p. 310). It seems to occur most often when 
the form of the electrodes is such as to drain the oil away by 
capillary forces. The only real remedy is to keep down the 
amount of lubricating oil going to the cylinders at starting and 
during glides. 

Cracking the insulator is one of the most common causes of 
spark plug failure. The thickness of the insulator is usually 
so great that a clear crack may not interfere with ignition, but 
after a while the cracks become filled with carbon and form a 
conducting path. 

Cracking may result from several causes. The high temper- 
ature gradient from the hotter inner end of the insulator to the 
relatively cold shell and the consequent unequal expansion is a 
frequent cause. Such cracks are most likely to occur at a 
shoulder or other place where there is a sudden change in diam- 
eter. Cracking may also occur if the metal parts of the plug are 
so arranged that their relatively greater expansion produces 
pressure on the insulator. The mechanical vibration of the 
engine as a whole may break the porcelain; such breakage often 
occurs in the outer portion of the porcelain. There is also 
considerable breakage from accidental mechanical injury such as 
striking the plug with a wrench. 

Mica plugs are free from this trouble. If porcelain is used it 
should combine high mechanical strength, low modulus of 
elasticity, low coefficient of thermal expansion and high thermal 
conductivity. The porcelain may also be made in two or more 
pieces-permitting the innermost porcelain to heat and expand 
considerably, while the outer pieces are cooler. The passage of 
a spark through the joint between the pieces is prevented by a 
21 



322 



THE AIRPLANE ENGINE 



wrapping of mica around both the shell and the electrode. It is 
very difficult to make such plugs gas-tight. 

In plugs in which the central electrode is cemented in the 
porcelain, the differential expansion of the two can be taken 
care of only if the electrode is kept of small diameter. 

Breakage by mechanical vibration can be reduced if the 
insulation is cushioned by a considerable thickness of asbestos 
or other packing material between the shoulder of the insulator 
and the bushing. One-piece plugs in which the edge of the shell 
is crimped over the shoulder of the insulator are especially liable 
to cracking at the edge of the shell. 




(«) (b) (c) (d) 

Fig. 257. — Types of spark plug construction. 



A minor cause of failure is the change in the width of the 
spark gap either through warping of the wires or corrosion. 
Warping occurs only when the wires are relatively long. Cor- 
rosion is very slow with the alloy commonly used (Ni 97 per 
cent, Mn 3 per cent), although a chemical reaction between the 
cement and the metal of the electrode may sometimes cause 
the tip to drop off. 

Gas leakage is an evil in that it causes a rapid heating of the 
plug if its amount is considerable; such leakage is usually a 
matter of workmanship rather than of design of the plug. There 
are two joints to keep tight, that between the central electrode 
and the insulator, and that between the insulator and the shell. 

The general methods of construction are shown in Fig. 257. 
In the screw bushing, a, the insulator has a shoulder, one side 
of which is seated on a shoulder in the shell while a bushing is 
screwed down inside the shell on the opposite side. A gasket 



IGNITION 



323 



of brass, copper, asbestos, or some soft heat-resisting material is 
used and can be placed on either side of the shoulder of the 
insulator, or on both. With a mica insulator it is possible to 
dispense with the gasket. To relieve the insulator from the 
mechanical strain resulting from differential expansion of the 
shell and the insulator such constructions as those shown 
diagrammatically in Fig. 258 may be used. In a the shell A and 
sleeve B are made of different metals and B is of such length as 
to maintain constant pressure on the washer. The expansion 
of the central electrode is compensated in a similar manner. 
In b the steel clamping nut is very thin and flexible. Expansion 
of the central electrode is provided for by a strong spring washer 
under the nut B. 





(a) (6) 

Fig. 258. — Special spark plug constructions. 

The crimped shell, b (Fig. 257), is most common (Champion, 
Titan, etc.) and is formed by forcing the top edge of the shell 
over a gasket, which rests on the upper side of the shoulder of 
the insulator. These plugs cannot be disassembled. 

The taper fit, c (Splitdorf), is used with a mica or steatite 
insulator. The mica does not stand well the pressure exerted 
on it during assembly. If a thin steel jacket is placed over the 
taper it protects the mica and is flexible enough to form a gas- 
tight fit with the shell. 

A molded-in insulator, d (Anderson), consists of glass which has 
been forced between the central electrode and the shell while in 
the molten state. It adheres to both electrode and shell and is 
gas-tight. 

The shape and arrangement of the electrodes seem to have 
little effect on the operation with the exceptions already noted 
of the greater liability to fouling with oil of plugs in which 



324 THE AIRPLANE ENGINE 

the side wall of the shell forms one of the electrodes. The 
variation in breakdown voltage with the shape of the tips is 
slight. With fine wires any oil film at starting is burned off 
rapidly but there is greater liability to preignition. With a 
central electrode consisting of a disc (Fig. 257c) the danger of 
short-circuiting with carbon is increased while the likelihood 
of complete fouling with oil is diminished and greater protection 
is afforded to the insulating material back of it. 

The location and number of spark plugs used are important 
in their effect on engine performance. An effort should be made 
to reduce as much as possible the distance through which the 
flame has to be propagated. The greater that distance, the 
greater is the liability to detonation. If the distance is shortened, 
a higher compression may be employed without detonation. 
Where a single plug is used its location should be in the center 
of the head. Multiple spark plugs are desirable, not only to 
insure ignition in case of failure of one plug but also because they 
permit higher compression pressures. With a constant com- 
pression pressure the power is increased by increasing the number 
of plugs; for example, in a 5}4 by 6^-in. four-valve single- 
cylinder test engine with a compression ratio of 5.4 and at 
1,800 r.p.m., the brake horse power was increased 5.7 per cent 
by the use of two spark plugs and 11.1 per cent by the use of 
four plugs. 

A typical wiring diagram is shown in Fig. 259, which shows 
the wiring of a 12-cylinder Vee engine equipped with two mag- 
netos for regular operation, and a starting magneto. The 
regular magnetos also have radio connection. 

The relative advantages of battery and magneto ignition 
have been much debated. The performance of the engine is 
not affected by the source of primary current. With battery 
ignition the engine is started by turning it with the current on 
but fully retarded. With magneto ignition the engine is pulled 
over a few turns with no current on so as to fill the cylinder with 
an explosive mixture, and a starting magneto is then operated 
by hand, giving a shower of sparks in one cylinder; the starting 
magneto is arranged to be fully retarded. The element of 
danger in starting the engine is eliminated by this latter method 
of operation. The battery requires more attention than the 
magneto, particularly if the engine is to stand idle for a while; 
if it runs down there remains no means for starting the engine. 



IGNITION 



325 



The battery system is also more complicated than the magneto 
system but it lends itself better to irregular explosion intervals; 
the magneto must be of special type in this case (see p. 300). 
The generator in the battery system requires attention to keep 
commutator and brushes in good condition; there is no corre- 
sponding attention necessary with the magneto. The total 
weight of the magneto system, including dual magnetos and a 
starting magneto, is somewhat greater than that of a battery 



Right Distributor --**: 
Left Distr,buto>~ ( m - 



PropellerEnd 



Right Distributor 
Left Distributor 




Radio Connection 



Right Magneto 



Fig. 259. — Typical wiring system for 12-cylinder Vee engine. 

system. A magneto system is easier for the pilot to operate; 
there is only one switch handle to control and no ammeter to be 
watched. The battery system has distinct advantage when 
current is required for an electric starter, lights and other uses. 
The firing order adopted in actual engines is given below. 



8-cylinder 90-deg. Vee ' 



1L, 1R, 2L, 2R, 4L, 4R, 3L, 3R (Curtis 0X5, VX; 

Sunbeam Arab) 

1L, 4R, 2L, 3R, 4L, 1R, 3L, 2R (Hispano-Suiza) 

(counting from propeller) 



12-cylinder 60-deg. Vee 



326 THE AIRPLANE ENGINE 

12-cylinder 45-deg. Vee: 1L, 6R, 5L, 2R, 3L, 4R, 6L, 1R, 2L, 5R, 4L, 3R 

(Liberty) (counting toward propeller). 
1L, 2R, 5L. 4R, 3L, 1R, 6L,5R, 2L, 3R, 4L, 6R 
(Rolls-Royce Falcon and Eagle) 
1L, 1R, 5L, 5R, 3L, 3R, 6L, 6R, 2L, 2R, 4L, 4R 
(Sunbeam Maori, Cossack) 
9-cylinder Rotary: 1, 3, 5, 7, 9, 2, 4, 6, 8 (Gnome, LeRhone, BR1, BR2, 

Clerget) 
6-cylinder Vertical: 1, 5, 3, 6, 2, 4(Beardmore, Galloway, Siddeley, Austro- 

Daimler, Mercedes, Fiat) 
7-cylinder Radial: 1, 3, 5, 7, 2, 4, 6 (A. B. C. Wasp) 
9-cylinder Radial: 1, 3, 5, 7, 9, 2, 4, 6, 8 (A. B. C. Dragonfly) 



CHAPTER XIII 
LUBRICATION 

All rubbing surfaces in an engine should be lubricated. The 
most important of these surfaces are the cylinder walls, main 
bearings, crankpins, piston pins, and camshaft bearings, but 
there are also numerous other parts to be lubricated. 

The coefficient of friction of a bearing with good lubrication, 
moderate pressures and high speeds is practically independent of 
the materials composing the rubbing surfaces, but is proportional 
to the viscosity of the oil, to the rubbing speed and to the area; 
it is independent of the pressure. For high pressures and low 
speeds these laws do not hold; for velocities from 100 to 500 ft. 
per minute the coefficient decreases about as the square root of 
the velocity, for velocities from 500 to 1,600 ft. per minute it 
decreases about as the fifth root of the velocities, while above 
1,600 ft. per minute it is practically constant. With high pres- 
sures the coefficient of friction increases. 

In an airplane engine the loads on the principal bearing surfaces 
are variable, going through a cycle of changes every two revolu- 
tions of the engine, and varying from a maximum to a low mini- 
mum. For example, in the Liberty engine the total force on the 
crankpin varies from 4,980 to 1,500 lb.; on the intermediate main 
bearing from 7,250 lb. to 800 lb.; on the end main bearing from 
4,025 lb. to zero; on the center main bearing from 7,700 lb. to 
2,500 lb.; the piston side thrust from 930 lb. to zero. Further- 
more, the direction of the force changes in these principal bearing 
surfaces, so that the portion of the bearing which at one instant 
is supporting maximum pressure is later relieved of all pressure. 
This intermittent application of the load is favorable to good 
lubrication and permits the use of maximum pressures greatly 
in excess of what would be possible with continuous loading. 
The oil film which is squeezed out by the application of the maxi- 
mum pressure is replaced during the reduction or reversal of the 
pressure. 

The maximum load per square inch of projected area is great- 
est on the piston pin, which has a diameter considerably less than 

327 



328 THE AIRPLANE ENGINE 

the crankpin; in the Liberty engine this is 2,580 lb. per square 
inch as against 932 lb. on the crankpin; 1,675 lb. on the center 
main bearings; 1,580 lb. on the intermediate main bearings; 
815 lb. per square inch on the end main bearings. 

The rubbing speeds of the main bearings of airplane engines 
range usually from 16 to 20 ft. per second; the rubbing speeds at 
the crankpins will be somewhat lower. 

The total friction work at any bearing is proportional to the 
product of the mean total load by the rubbing speed. The limit- 
ing factors for a bearing for continuous operation are the mean 
pressure per square inch of projected area and the rubbing speed. 
The product of these two is a good index of the service of the 
bearing. In the Liberty engine this load index is 13,500 lb. -ft. 
sec. for the crankpin; 24,670 for the center main bearing; 13,650 
for the intermediate main bearing; and 11,900 for the end main 
bearings. 

The permissible pressure on the bearing depends on the viscos- 
ity and therefore on the temperature of the oil. The temperature 
tends to rise and must be kept down by oil cooling. Tempera- 
tures of 160°F. and higher are common. 

The lubrication of the cylinder offers problems quite different 
from the lubrication of the rest of the engine. The side thrust 
pressures are moderate; the piston speed is high, reaching 2,000 ft. 
per minute, or 33 ft. per second, and the maximum speed (at 
mid-stroke) is about 52 ft. per second. The friction work under 
these conditions would not be a serious charge against the engine 
if the oil film could be maintained in good condition, but as pointed 
out on page 24 the viscosity of the oil film on the cylinder walls 
is greatly raised by carbonization. The oil film on the walls 
and around the piston rings has to serve another purpose besides 
acting as a lubricant; it acts as a seal to prevent the blowing of the 
gases past the piston. For this purpose high viscosity is useful. 

In starting cold, in idling with overrich mixtures, and in cold 
weather, a certain amount of liquid fuel will meet the cylinder 
walls and, being perfectly miscible with the mineral lubricating 
oil, it will dilute the oil film and the thinned oil will then run down 
the cylinder walls and dilute the oil in the crankcase. This 
phenomenon, which is very common in automobile practice, is 
not so usual in airplane engines because of the higher volatility 
of the aviation fuels. The ordinary gasoline of commerce has a 
high end point, which means that it has kerosene constituents 



LUBRICATION 329 

which will not vaporize during admission and will be deposited 
in the liquid form on the cylinder walls. With aviation fuel the 
dilution when it occurs is by more volatile elements which tend 
to vaporize out from the hot body of oil so that the dilution of 
the crankcase oil is not cumulative as with automobile engines. 

The friction at the bearings of an engine results in heat which 
has to be taken away as fast as it is generated if the bearings are 
not to rise in temperature. Much of this heat is conducted 
through the metal to cooler parts of the engine but part is carried 
away by the oil itself. For this purpose a large flow of oil is 
desirable. The amount of oil circulated in the Liberty engine is 
about 12 gal. per minute and the temperature rise may average 
about 10°F. This corresponds to a heat abstraction of about 
45 B.t.u. per minute. If the bearing friction work is taken as 
1J^ lb. per square inch of piston area (see p. 24), the correspond- 
ing heat generated will be about 200 B.t.u. per minute. The 
volume of oil circulated in this case is not sufficient for complete 
cooling of the bearings. 

Viscosity. — The oil which gives minimum temperature rise of 
the bearing is the best to use, other factors being equal. An oil 
of lower viscosity will cause greater friction because it will squeeze 
out and allow a closer contact of metallic parts and increase 
metallic friction and wear. An oil of higher viscosity results in 
increased fluid friction of the oil. With a complete film of oil, 
the oil flows like a pack of playing cards sliding over each other, 
the outer layers adhering to the surfaces and not sliding with 
reference to them. The actual fluid friction F, in pounds, is given 
by the equation 

PXA X V 
5,760 X t 

where P is the absolute viscosity in poises; A is the rubbing area 
in square inches; V is the rubbing velocity in feet per second; 
and t is the oil film thickness in inches. This formula indi- 
cates that the friction diminishes as the thickness of the film 
increases. 

The measurement of viscosity has been standardized in this 
country and is determined by the Saybolt Universal Viscosimeter 
in which the fluid to be analyzed flows through a tube 0.1765 cm. 
diameter, 1.225 cm. long, under an average head of 7.36 cm., from 
a vessel 2.975 cm. diameter. The time in seconds required for 
60 c.c. of oil to flow through the tube is the viscosity in seconds 



330 THE AIRPLANE ENGINE 

Saybolt. The absolute viscosity is obtained from the equation 

180n 



100 



P =G(0.22S-^) 



where G is the specific gravity of the oil and $ is the Saybolt 
viscosity. 

One of the most troublesome features of lubrication is the 
decrease in viscosity of oils with rise in temperature. It is found 
that if the logarithm of the Saybolt viscosity is plotted against 
the temperature (Farenheit), on cross-section paper, the points 
lie close to a straight line — this is a purely accidental relation. 

The viscosities of the principal American lubricating oils at 
temperatures of 100°, 150° and 212°F. are given in Table 16. 
It will be noticed that the very heavy oils fall off most rapidly in 
viscosity as the temperature is raised so that whereas the range 
in Saybolt viscosities of the oils tabulated at 100°F. is about 11 to 
1, at 212° it is only 3 to 1. 

The desired physical characteristics of a lubricating oil are as 
follows : 

(a) Body sufficient to prevent metallic contact under maximum pressure 
and maximum temperature. 

(b) Lowest viscosity in keeping with the above conditions. 

(c) Capacity of resisting high temperatures without decomposition. 

(d) Fluidity at minimum temperatures. 

(e) High fire test. 

(J) Freedom from oxidation. 

(g) Freedom from corrosive action on metals. 

The standard specifications for lubricating oils for airplane 
engines adopted by the U. S. Army and Navy are given below. 
Grade 1 is the Navy specification, Grade 2 the Army. The 
specifications are also different for summer and winter use. 

Flash Point. — The flash point of Grade 1 shall not be lower 
than 400°F.; for Grade 2 not lower than 500°F. 

Viscosity at 210°F. shall be within the following limits: 

Grade 1 (summer) 90-100 sec. 

Grade 1 (winter) 78- 85 sec. 

Grade 2 125-135 sec. 

Pour Test. — Grade 1 not above 45°F. for summer, or 15°F. 
for winter. 

Cold Test.— Grade 2 not above 35°F. 



LUBRICATION 



331 



Table 16. — Properties of Representative American Lubricating Oils 
for Use in Internal Combustion Engines 



Physical properties 



Kind of oil 



Baume 
grav. 




Viscosity, seconds 



Deg.F.J Deg. F. , 
burn i chill 




Havoline: 
Light.... 
Medium . 



25.9 
25.0 
Heavy 25.6 

Mobiloil: 

"E" Light 28.1 

"A" Medium '< 21.8 

"B" Heavy 26.3 

Arctic Lt. Med 23.3 

Arctic Medium 21.1 

"BB" Med. Heavy 25.8 



Monogram: 

Light 27.6 i 360 

Medium 26.0 375 

Heavy 28.9 430 

Ex. Heavy 24.7 465 



Perfection: 
"A" Light..., 
"B" Medium. 
"C" Heavy. . 



Socony: 

Zero 

Polarine Heavy. 



Texaco: 
Light.... 
Medium. 
Heavy. . 



29.1 
24.9 
29.3 



24.3 
25.4 



21.3 
20.9 
19.3 



26.2 
27.1 
26.3 



Veedol: 

Aero No. 1 

Aero No. 2 

Aero No. 3 

Aero No. 4 27.6 

Aero No. 5 24.7 

Aero No. 6 24.7 

26.2 
26.1 



Zero Heavy 

Zero Extra Heavy 



Wolf's Head: 

Heavy 28.6 

No. 8 ! 27.6 

Castor oil 15.0 



370 
385 
395 



370 i 

360 ! 

500 

370 

370 

460 



400 
390 
420 



395 
385 



335 
350 
356 



455 
450 
435 
440 
440 
460 
410 
465 



415 
485 



380 
395 
410 



380 
360 
470 
380 
585 



430 
450 
455 



420 
420 
580 
425 

430 
540 



360 410 

370 430 

445 505 

425 535 



410 
400 
430 



410 
380 



340 
350 
360 



450 
445 
435 
430 
450 
460 



460 



470 
450 



535 
530 
520 
515 
520 
540 
480 
550 



475 
550 





24 
41 



46 



470 26 
450 32 
495 40 



SI at 
35 



380 SI at 
400 SI at 
420 10 



173 
237 
361 



181 
243 
316 



219 
300 



205 
301 

495 



66 

80 

111 



167 66 

330i 97 

1,640; 397 

221 | 74 

300, 87 

926! 243 



140 60 

289 95 

340 108 

1,583! 356 



71 

81 

103 



77 
103 



85 
119 



795 


212 


814 


222 


517 


149 


513 


151 


413 


135 


474 


134 


329 


107 




355 



334 108 
1,196 300 
1,2701 305 



42 
46 
54 



44 
49 
122 
45 
46 
86 



41 

50 

55 

110 



78 
80 
63 
64 
55 
58 
52 
111 



52 

100 

90 



332 THE AIRPLANE ENGINE 

Acidity. — Not more than 0.10 mg. of potassium hydroxide 
shall be required to neutralize 1 gram of Grade 1 oil. 

Emulsifying Properties. — The oil shall separate completely 
in 1 hr. from an emulsion with distilled water at a temperature of 
180°F. 

Carbon Residue in Grade 1 shall not be over 1.5 per cent; 
in Grade 2 not over 2.0 per cent. 

Precipitation Test. — When 5 c.c. of the oil is mixed with 
95 c.c. of petroleum ether and allowed to stand for 24 hr. it 
shall not show a precipitate or sediment of more than 0.25 c.c. 

The oil shall not contain moisture, sulphonates, soap, resin 
or tarry constituents. 

The Flash Test shows the temperature at which the vapor from a sample, 
heated in an open cup, will ignite. It has some relation to loss by evapora- 
tion. The open cup is 23^ in. diameter, l^{& in. high and is filled to within 
% in. of the top. It is placed on a metal plate and heated so that its tem- 
perature rises not less than 9° nor more than 11°F. per minute. A test 
flame %2 in. in diameter is passed across the top of the cup, taking 1 sec. for 
the passage, at every 5°F. rise of temperature of the oil. The flash point is 
the temperature at which a flash appears at any point on the surface of the 
oil. Drafts must be avoided. 

The Viscosity Test has been discussed on page 329. 

The Pour Test indicates the temperature at which a sample of the oil 
will just flow. The oil is placed in a glass jar 134 in. diameter and 4 to 5 in. 
long to a depth of about 34 in. and the jar is corked. It is then placed in a 
freezing solution and at each 5°F. drop in temperature it is taken out and 
tilted. The pour test is 5° higher than the temperature at which the oil will 
not flow when the jar is placed in the horizontal position. 

The Cold Test has a similar purpose to the Pour Test but in this case the 
oil is first frozen and is then stirred with a thermometer until it will run from 
one end of an ordinary 4-oz. sample bottle to the other. The temperature 
reading at that time is the Cold Test. 

Carbon residue is obtained by heating 10 grams of oil in a porcelain crucible 
placed inside two iron crucibles with covers. It is heated so as to maintain 
a vapor flame of specified length and heated further after the vapors cease 
to come off. After cooling the weight of carbon residue in the porcelain 
crucible is determined. 

Reclaiming Oil. — The oil used in a stationary airplane engine 
not only becomes diluted by the heavier constituents of the fuel 
but also becomes dirty by the accumulation of free carbon from 
the cylinder walls, of metal particles worn off the bearing surfaces, 
and of other solid impurities. The oil usually has to be changed 
between the fifth and twentieth hour of flying service; most oil 
is not used more than 5 hr. It can be reclaimed by allowing it to 



LUBRICATION 333 

stand 30 hr. in a tall bucket, decanting off the upper two-thirds, 
filtering and warming to 150°F. A more thorough process is to 
put the oil with some water and soda ash in a steam-jacketed 
tank, raising the temperature to 212°F., forming an emulsion 
and obtaining a precipitation of carbon, iron, and dirt after a 
period of rest. The steam drives off the 2 or 3 per cent of diluent 
coming from the volatile aviation gasoline. A recovery of 85 
per cent is possible in this way and the reclaimed oil is at least 
as good as new oil. 

A centrifugal oil cleaner has been tried on the Liberty engine 
with considerable success. This consists of a spun copper bowl, 
5 in. diameter, rotating at lj^ crankshaft speed; the centrifugal 
force at 2,550 r.p.m. is about 45 times that of gravity. The oil is 
led into the center of the bowl and is thrown out at the top. 
Examination of the contents of the bowl after a run show that it 
collects metal particles, sand, carbon, rubber and other solids. 
Its use should increase considerably the periods between changes 
of oil and should prolong the life of all bearing surfaces by 
preventing their abrasion by solid particles in the oil. 

Castor oil is employed in rotary engines in which the gasoline is 
admitted to the crankcase on its way to the cylinders. Mineral 
oil (petroleum) cannot be used in this case as it is miscible with 
gasoline and its use would result in a thinning of the lubricant 
and a wastage of fuel. 

Methods of Lubrication. — The splash system of lubrication 
often employed in automobile engines is not satisfactory for 
airplane engines on account of the high loading at which they are 
operated and also because of the extreme variations in engine 
orientation during flight. For the last reason also a wet sump is 
undesirable since it will deluge some of the cylinders during such 
airplane evolutions as a nose dive and may result in trouble from 
excess of oil in the cylinder. Consequently the modern airplane 
engine is provided with a pressure oiling system and a sump, 
which is usually kept dry by a scavenger pump. 

The normal lubricating system is as follows. The scavenger 
pump or pumps take oil from the sump and deliver it to the 
external oil tank. The pressure pump takes oil from the oil 
tank and discharges it into a distributing main from which 
branches go to each of the main bearings. Oil enters some of 
the hollow main journals through small holes which register 
with corresponding holes or channels in the bearing and there- 



334 THE AIRPLANE ENGINE 

by with the branch oil pipes. Usually, alternate journals are 
rilled with oil and the crank cheeks on the two sides of these 
journals are drilled to connect with the hollow crankpins and 
thereby permit lubrication of the crankpins through appropriate 
holes in them. The oil-containing journals and all the crankpins 
have closed ends (see p. 144). The lubrication of the piston pin is 
sometimes carried out by oil pipes running along the connecting 
rod and registering every revolution with the oil hole in the crank- 
pin; in other cases the oil thrown out by centrifugal force from the 
crankpin is relied on to lubricate the piston pin as well as the 
cylinder wall. 

The camshaft is lubricated from an oil pipe from the end 
of the distributing main, connecting with it usually by an annular 
groove in the front main bearing. The camshaft is hollow and 
acts as an oil carrier discharging oil through small holes at each 
bearing. The oil escaping from the bearings lubricates the cams 
and returns to the crankcase over the distributing gears, meeting 
there the oil escaping from the main bearings, crankpins and 
cylinders. 

The lubricating system of the Liberty engine follows the lines 
indicated above. The cylinders, pistons and wristpins are 
lubricated by oil spray from the crankpin. A double scavenging 
pump at the rear of the engine (Fig. 261) keeps the two sumps 
drained and returns the oil to the outside oil tank. In the 
Hispano-Suiza engine (Fig. 51) a dry sump is also used. A 
single gear-type scavenging pump is driven directly from the 
rear of the crankshaft; an eccentric sliding-vane pressure pump 
is mounted on the vertical water-pump shaft and rotates at 1.2 
times the crankshaft speed. The vane pump forces the oil 
through a filter to the main oil pipe in the lower crankcase. 
There are four oil holes in each crank pin through which oil goes 
to the bearing and is thrown off to lubricate the cylinder and 
wrist pin. A small hole is provided in the leading face of each 
cam to lubricate the cam and its follower. 

In the Curtiss K.-12 (Fig. 55) the pumps are located in the 
lower part of the crankcase and are driven through a horizontal 
shaft. There is a triple-gear scavenging pump and two pressure 
pumps arranged in a unit with the spiral driving gears and 
surrounded by a filtering screen. The oil is supplied to the main 
bearings in the usual way and is then conveyed to the crankpins 
through small tubes built into the crankshaft. The oil is cooled 



LUBRICATION 335 

by a temperature regulator through which the jacket water 
circulates on its way from the radiator to the pump. 

The Curtiss OX engine uses a wet sump (Fig. 59) covered 
by an oil-pan partition. A single-gear pump driven from a beveled 
gear on the crankshaft at the propeller end sucks oil from the 
sump and discharges it into the rear end of the hollow camshaft 
whence it goes through tubes to the crankshaft bearings. In this 
engine there is a continuous closed oil passage from one end of the 
crankshaft to the other. The cylinder is lubricated by oil spray. 
The oil-pan partition has a half-inch hole at its center for the 
return of oil to the sump. 

The Hall-Scott L-6 engine (Fig. 63) uses a wet sump and has 
scavenger and pressure pumps mounted as a unit inside the lower 
crankcase and driven through an inclined shaft from a bevel gear 
which is at the rear of the crankshaft. The system is otherwise 
similar to the Liberty engine. An oil sight gage (Fig. 64) 
shows the level in the sump. Splash plates in the lower case pre- 
vent excessive splash from the dipper action of the connecting 
rods. 

In the Napier "Lion" engine three oil pumps are combined as 
a single unit at the extreme rear of the engine ; they are of the gear 
type and are driven at half engine speed. The suction pumps draw 
the oil away from the two ends of the lower crankcase by two 
separate steel pipes (Fig. 72) and discharge to the tank through 
a common pipe. The pressure pump delivers to both ends of the 
crankshaft and to the three camshaft casings. There is a con- 
tinuous passage for oil through the crankshaft. 

In the Fiat -650 engine (Fig. 76) there are scavenger pumps at 
the two ends of the lower crankcase and a pressure pump di- 
rectly under the rear scavenger pump. They are driven by bevel 
gears from the horizontal tubular shaft inside the crank casing 
and are mounted in ball bearings as well as the horizontal shaft 
and the spur gear which drives it. The oil drawn from the main 
tank is discharged into a copper main cast into the crankcase. 
The main bearings are fed from copper branch pipes. In other 
respects the system is normal. 

In the Benz-230 (Fig. 77) a wet sump is used and the triple 
oil pump is submerged in the sump. The main pressure pump A 
(Fig. 260) draws oil from the reservoir in the sump and discharges 
it to the main bearings through a distributing main and branch 
pipes. The supply to the piston pins is through small pipes 



336 



THE AIRPLANE ENGINE 



inside the tubular connecting rods. Fresh oil is fed into the 
sump by the small suction pump, B, from the oil tank while 
the correct working oil level is maintained in the reservoir by the 
pump, C, whose curved suction pipe (see Fig. 77) terminates at 
the desired oil level. All return oil passes over the transverse 
air pipes in the lower crankcase and is cooled by them. 

In the Maybach engine (Fig. 80) scavenger pumps are 
mounted at both ends of the crankcase and the pressure pump is 
placed behind the rear scavenger pump. All three are operated 
by the same horizontal shaft driven by spur gearing from the 
front end of the engine shaft. The oil is discharged into an 



Return Pipe to 

Tanktrom 

PumpC 



Main Delivery Pipe 
from_ Tank to Pump B 



Oil Level in Sump 




From Sump 
From Sump To Tank / n to Pump-C 

b / a / cf 



FromSur, 



Delivery to 
Main Bearings 



A- Pump Supplying B-PumpSupplyinq, OPump 
Main Bearing * Sump from Returning 

From Sump OH Tank Oil from Sump 

To Tank 



From Sump Supply to 
toPump-A Sump from 
Pump-B 
D- End View 
of Pump 
Cover. 



Fig. 260. — Triple-gear pump of Benz engine. 



external oil main on the upper crankcase and past individual 
screens into branch pipes drilled through the transverse webs 
into the main bearings. The oil thrown off from these bearings 
is caught in aluminum scoops bolted to both ends of each crank- 
pin, is carried by centrifugal force into the hollow crankpin and 
thence through radially bored holes to the bearing. The piston 
pins are oiled through the internal pipes in the connecting rods. 
Baffle plates bolted to the upper crankcase just below the cylin- 
ders prevent an excess of oil reaching the cylinders. 

Relief Valves. — All pressure-feed systems are provided with a 
relief valve on the discharge side of the pump. This is a spring- 
loaded valve as in Fig. 261 and is set for the maximum allowable 
pressure. The oil pressure is high in starting especially in cold 



LUBRICATION 337 

weather when the viscosity of the oil is very great. The normal 
operating oil pressures after fully warming up are about 25 to 30 
lb. per square inch in Liberty engines, 50 to 60 lb. in Curtiss 
engines, 40 to 65 lb. in Hispano-Suiza engines. In cold weather 
it is best to drain off the oil after a flight and to fill up with 
hot oil before starting. 

The location of oil grooves in the bearings is a matter of con- 
siderable importance on which there is much divergence of 
practice. The actual pressure on the oil film will vary from zero 
at the ends of the bearings and at the split of the bearing to 
possibly as much as 10,000 lb. per square inch in the center of the 
loaded area at the moment of maximum loading. As the oil 
pressure does not exceed 50 lb. per square inch it is obvious that 
oil cannot be forced in at the place of maximum loading unless the 
pressure at that place falls below 50 lb. per square inch during 
some part of the cycle. There should be no oil grooving length- 
wise in the middle of the most loaded half of a bearing; such 
grooves are channels of escape for the oil and may result in such 
thinning of the film as to increase the friction and, possibly, to 
cause seizing. The most heavily loaded part of the main bearing 
is usually the middle of the lower cap and it is at this place that 
the oil usually enters. The grooves should then be two helical 
grooves intersecting at the oil hole at the center of the bottom 
half of the bearing and running to the split but not too near the 
ends of the bearing. Short helical grooves in the upper half of 
the bearing may start at the split opposite the lower grooves but 
should not go more than half-way up each side. Similar groov- 
ing should be provided at the crankpin bearing, which also is most 
heavily loaded at the lower half. 

Wherever practicable it is desirable that the oil should enter at 
the place of minimum average bearing pressure. It is the 
cyclical variation in the loading that makes possible the proper 
lubrication of the heavily loaded bearing surfaces of aviation 
engines. The oil pressures employed are in themselves not nearly 
adequate to support the loads but are required to overcome the 
viscous and frictional resistance to the flow of the oil to the va- 
rious bearings and also to ensure that the oil channels will clear 
themselves of small obstructions. 

The amount of oil circulated per minute is determined not 
only by the lubrication needs but also, as pointed out on page 329, 
by the extent to which the oil is used as a cooling medium. For 

22 



338 THE AIRPLANE ENGINE 

lubrication the amount should probably be some function of the 
total projected areas of the main bearings and crankpins; 
expressed in this way the use of oil varies from 0.1 to 0.5 lb. per 
square inch of projected bearing area per minute. In terms of 
the power delivered by the engine the oil circulated varies from 
0.025 to 0.15 lb. per horse-power minute. 

The oil consumption of an engine as usually measured is the 
amount of oil which has to be added to the system to make up for 
oil burned or otherwise used up during the engine operation. 
This quantity varies from 0.02 to 0.05 lb. per horse-power hour 
in stationary water-cooled engines but may go as high as 0.15 
lb. in rotaries. 

Oil Pumps. — The great majority of airplane engines use gear 
pumps both for scavenging and pressure pumps. A simple gear 
pump consists of a power-driven spur gear meshing closely into an 
exactly similar driven gear, the gears being enclosed in a casing 
with the minimum working clearance above and below the gears 
and also around them except where they mesh. The oil inlet is on 
the side where the gears separate; the discharge is on the side 
where the gears meet. If there were no leakage the oil carried 
from the inlet to the discharge side would be equal to the space 
between the teeth, but some of this is brought back to the inlet 
side since a tooth going into mesh does not fill the space between 
adjacent teeth and consequently does not displace all the oil 
content of that space. The capacity of each gear can be taken 
approximately as equal to half the annular space between the 
roots and tips of the gear, or, for the two gears, as equal to the 
whole of the annular space of one gear. As the width of this 
annular space increases with decrease of the number of teeth 
(decrease of pitch) it is evident that, for a given pitch diameter, 
capacity can be increased by decreasing the pitch. 

Two scavenger gear pumps are sometimes combined into a 
triple-gear pump as in the Liberty engine. In this case the 
driving gear is central and the inlets to the driven gears are on 
opposite sides of the pump. 

The driving gears for scavenger and pressure pumps are usually 
placed close to one another and driven by the same shaft. In the 
Benz engine (Fig. 260) three such gear pumps are mounted on the 
driving shaft and function as indicated in the figure. The 
Liberty oil pump is driven at one and one-balf times crankshaft 
speed and has a capacity of 1.9 gal. per minute at normal speed. 



LUBRICATION 



339 



The pump consists of a double scavenging pump with three 
gears, A, B and C, Fig. 261, drawing oil from the two sumps and 




yOears in Upper 
m Pocket 



.Oil from Lower 
Pocket Under 
Pressure 



Oil from Tank 



Dram Plucf-'-^ 



Cross-sec+ion of Liberfu-12 Oil Pump 
Fig. 261. — Gear pump of Liberty engine. 

a pressure pump immediately below it giving an oil pressure of 
35 to 50 lb. per square inch at engine speeds from 1,500 to 1,800 
r.p.m. The pressure gears, A r , C, (Fig. 261) are immediately 



340 



THE AIRPLANE ENGINE 



below the gears A and C of the scavenging pump. The gear 
A' is driven from the pump shaft and the upper train is operated 
through a vertical-shaft connection from C to C. The pressure 
pump draws oil from the tank through a copper pipe, the oil 
passing through the large lower strainer and entering the gear 
housing at M; it is discharged through the passage NOP to the 
distributing manifold. The pressure relief valve is shown in 




1400 1600 

Rev per Min of Engine 

Fig. 262. — Performance curves of gear pump of Packard-180. 



Fig. 261; the spring is usually set for a pressure of 50 lb. per 
square inch. The discharge from the relief valve goes to the 
suction side of the pressure pump. 

The oil from the rear sump entering the scavenging pump after 
passing through the upper strainer is drawn through E and dis- 
charged to the outlets F and G, which are connected by a passage 
in the pump body. From G, the oil goes through the passages K 




Fig. 263. — Sliding-vane pump. 

and L to the oil tank. Oil from the front sump is taken through 
an internal pipe and the passage HIJ and is also discharged at 
F and G. 

The volumetric efficiency of the gear pump can be made very 
high — probably up to 90 per cent; the over-all efficiency is 50 to 60 
per cent. Tests of the pressure pump of the Packard 180-h.p. 
engine give the results shown in Fig. 262. It will be seen that 
the slip is very low since the discharge curves are almost the 



LUBRICATION 



341 




From Sump 
Fig. 264. — Plunger pump of Basse-Selve engine. 




Main Pressure Pumps A 

Scavenger Pumps & 

Delivery of Pumps A «^^^ 
Suction of Pumps A ■■ " ■■ 
Delivery of Pumps B 
Suction of Pumps B 
Gravity Feed from Oil Tank 

Oil Radiator — C 

Pulsation Damper D 



Fig. 265. — Oil system of Basse-Selve engine. 



342 



THE AIRPLANE ENGINE 



same with free discharge and with discharge against pressures 
which vary from 48 lb. at 1,200 r.p.m. to 58 lb. at 2,000 r.p.m. 
It may be further noted that the slopes of the discharge curves up to 
1,600 r.p.m. are such as to go through the zero of ordinates; 
that is, the pump discharge is directly proportional to its r.p.m. 
The eccentric sliding-vane pump (Fig. 263) is used occasionally 
but is being displaced by the gear type. The sliding vanes are 




To Delivery 

Main. .. 




Fig. 266. — Plunger pumps of Salmson engine. 

pressed [out by springs in a slotted cylinder which is mounted 
eccentrically in a cylindrical casing and is power-driven. 

Plunger pumps are used in some of the German engines, being 
operated by eccentrics or by scroll cams. In the Basse-Selve 
engine, (Fig. 264) the oil pump is driven by a worm gear. The 
pump consists of two double-acting steel plungers which work 
vertically in aluminum barrels and make in effect four pumps. 




Fig. 267. — Action of LeRhone oil pump. 

The plungers are rotated by the worm gear and are simultan- 
eously reciprocated by the action of the scroll cam cut in the 
spindle and operated by a hardened steel roller working on a pin 
screwed into the pump body. At each stroke of the two plungers 
oil is drawn from the cooler tank to one of the inner pump cham- 
bers and is discharged from the other inner pump chamber into 
the delivery main. At the same time oil is sucked out of one of 
the engine sumps into one of the outer pump chambers and is 



LUBRICATION 



343 



pumped from the other outer pump chamber into the cooler 
tank. A diagrammatic view of the system is shown in Fig. 265. 
An entirely different arrangement of plunger pumps is used 
in the Salmson engine (Fig. 266). In this case two plungers are 
used, the larger one being the scavenging pump. The plungers 
are pivoted on a crankpin rotated by worm gearing. The pump 
bodies are pivoted and oscillate through a small angle as the 
crankpin rotates and the plungers make their strokes. The 
connections to suction and discharge are made when openings 
in the oscillating pump bodies register with appropriate openings 




End View Side View 

Fig. 268. — Mounting of oil tank. 

in the pump casing. The method of action of a single pump of 
this type, used on the Le Rhone engine, is shown diagrammatically 
in Fig. 267. The port, P, in the oscillating cylinder comes 
alternately opposite the intake port, /, and the delivery port, D. 
The method of mounting the oil tank and of connecting it to 
the engine in the USD-9A airplane is shown in Fig. 268. The 
tank is slung from the engine sills, L, and is located just below the 
crank case, C. The tube, A , is the oil return pipe from the scaven- 
ger pump to the tank; the supply pipe from the tank to the 
pressure pump is shown in dotted lines. The front sump con- 
nects by a pipe, B, to a Y-shaped air-pressure-relief pipe inside 
the tank, the upper ends of which extend nearly to the top of the 
tank. This acts as a pressure-relief vent for the tank. 
The oil is cooled by longitudinal air pipes, H. 



CHAPTER XIV 
THE COOLING SYSTEM 

All airplane engines are ultimately air-cooled. The only 
option is as to whether the air shall be applied directly or indi- 
rectly. In the latter case, the heat is removed by water which 
is then cooled by the air. Indirect (water) cooling offers two 
advantages: (1) the transmission of heat from the cylinder is 
more rapid to water than to air, and (2) the ultimate cooling 
surface (in the radiator) can be made much greater than is 
possible at the cylinder. Direct-air cooling has the advantage of 
reduced total weight and diminished vulnerability; a bullet hole 
through the radiator or jacket will put the whole engine out of 
action while a hole through one cylinder of an air-cooled engine 
may put that cylinder alone out of action. 

Air cooling in airplanes is used almost exclusively in rotary 
and radial engines. In vertical or Vee engines with several 
cylinders in line, the cooling problem is much more difficult, 
although it has been met by the use of suitable cowling to direct 
the air on to the different cylinders. With the rotary and radial 
types the motion of the plane and the location of the engine in the 
slip stream of the propeller ensure an adequate flow of air for 
cooling; with the multicylinder vertical or Vee type it is some- 
times necessary to add a fan to improve the air circulation. 

The resistance or drag of air-cooled cylinders is considerable 
but has not been determined experimentally in a satisfactory 
manner. In the rotary engine there is, in addition to the drag, 
the resistance due to churning which reduces directly the b.h.p. 
of the engine. This resistance increases so rapidly with speed 
that it is not found desirable to operate rotary engines at speeds 
in excess of about 1,400 r.p.m.; the increase in indicated power 
which results from increased speed is largely used up in over- 
coming the increased air resistance. The total work done by an 
air-cooled engine in overcoming air resistance is probably greater 
under ordinary conditions of operation than the total work done 
by a water-cooled engine in overcoming the drag of its radiator. 
Until quite recently, air-cooled cylinders were at a great dis- 
advantage both in fuel economy and in power developed per 

344 



THE COOLING SYSTEM 345 

unit volume of piston displacement, but recent constructions 
have put the air- and water-cooled engines very nearly on a par 
in these respects. They have always had an advantage in 
weight per horse power. 

The heat which has to be removed from the cylinder in order to 
keep the temperature of the cylinder within the limit which 
permits satisfactory operation of the engine is usually about 
equal to the heat equivalent of the work done in the cylinder, or 
42 B.t.u. per brake horse power per minute. This quantity 
will increase or decrease with change in operating conditions; 
its limits are apparently between 30 and 60 B.t.u. per brake 
horse power per minute. 

The cooling surface of the modern air-cooled cylinder almost 
invariably takes the form of a series of fins. Tests on the rate 
of heat dissipation from such surfaces, made by the British 
Advisory Committee for Aeronautics, 1 show that, for wind 
speeds between 20 and 60 miles per hour, the heat loss for a 
given material is independent of the roughness of the 
surface. Steel shows 5 to 10 per cent greater heat dissipation 
than aluminum or copper. Aluminum is improved about 10 
per cent by a coating of stove enamel. 

Throughout the usual range of cylinder diameters the heat 
dissipation for copper fins in a parallel air blast at the ground is 
given by 

H = [0.0247 - 0.0054(Z°- 8 /p°- 4 )]7 - 73 
where H is the heat dissipated in B.t.u. per square foot of fin 
surface per minute per degree Fahrenheit difference between the 
mean fin temperature and the incoming air temperature; I is the 
length of the fins in inches; p is the pitch in inches measured from 
surface to surface of adjacent fins; and V is the wind speed in 
miles per hour. With tapering fins p should be taken at the 
mean height of the fin. The heat dissipation, depending on the 
weight of air brought into contact with the cylinder, is 
proportional to (dV) - 73 , where d is the air density. 

Shape and Size of Fins. — The fin which gives the maximum 
heat-loss per unit of weight is one having slightly concave 
surfaces and a sharp tip ; a plain triangular fin is only very slightly 
less efficient. The best proportions for such a fin depend on the 
conductivity of the material and on the wind speed. For a 
speed of 40 miles per hour, the following table shows the best 

1 A. H. Gibson, Institution of Automobile Engineers of Great Britain, 1920. 



346 



THE AIRPLANE ENGINE 



proportions for fins of aluminum alloy (conductivity = 0.38 
C.G.S. units) and steel (conductivity = 0.12 C.G.S. units) and 
also for rectangular copper fins (conductivity = 0.90 C.G.S. 
units). 



Bottom breadth, centimeters 


0.025 


0.05 


0.1 


0.2 


0.3 


0.4 


0.5 


Length, centimeters: 

Aluminum 




2.3 


2.0 
1.1 
3.3 


2.9 
1.5 

4.8 


3.5 

1.8 


4.1 
2.1 


4.5 


Steel 




2.3 


Cooper 


1.6 









If such a fin be truncated until the tip breadth is one-fifth of 
the bottom breadth, the lengths become 80 per cent of those 
given above. The heat dissipation is about 0.88 time and the 
weight 0.96 time as great as for the complete triangular fin. 

Since the heat dissipated from a fin of given shape varies 
directly as the length of the fin, while the weight varies as the 
square of the length of the fin (other things being equal) cooling 
fins should be as short as possible, a large number of short thin 
fins being used in preference to a smaller number of longer and 
thicker fins. While in practice this is to be borne in mind, many 
other factors besides that of weight have an important bearing 
on the best size of fin to be adopted. Thus, in a thin steel 
cylinder, or in a cylinder of cast iron or cast-aluminum alloy, the 
circumferential ribs add greatly to the strength and resistance to 
distortion. Comparatively deep and heavy fins have a greater 
effect in this direction than a larger number of similar but smaller 
fins giving the same cooling. Again, as the number of fins in 
increased, the pitch is correspondingly diminished. This 
diminution in pitch reduces the air flow between the fins to an 
extent which may, with very small pitches, render the fins prac- 
tically useless for cooling purposes. 

In a cylinder of cast iron or of aluminum alloy, foundry 
difficulties put a definite limit to the minimum pitch of the fins. 
On the barrel itself a somewhat smaller pitch may be adopted 
than on the head, or the barrel fins may be turned out of the 
solid if desired. On account of the complicated form of the cylin- 
der head and ports, however, it is difficult to machine their 
cooling fins, and the length of many of the cores necessitates the 



THE COOLING SYSTEM 



347 



pitch being made fairly large. The minimum practical pitch 
of fin for such cylinders, having a diameter of from 4 in. to 6 in., 
is about 8 to 9 mm. or about %& in. Foundry difficulties also 
prevent the casting of a fin having a tip less than about 0.5 mm. 
in thickness, or a root thickness less than about Z/10, so that an 
aluminum fin 1 in. long would not have a root thickness less than 
0.1 in. 

For steel cylinders with fins turned out of the solid, the pitch 
may with advantage be cut down to about J4 in. on a cylinder of 
3 in. or so in diameter, but there appears to be little to be gained 
by reducing the pitch beyond this point. 

The mean fin temperature depends on the total amount of heat 
which has to be dissipated and its value depends on many factors. 
Determinations of the actual temperatures of the cylinder 
walls, pistons, and exhaust valves show no definite differences 
between well designed air-cooled and water-cooled engines; the 
air-cooled cylinder may be, and often is, the cooler of the two. 

The influence of engine speed on the wall temperature is 
shown in the following table giving the temperature at a point on 
the side of the combustion space of an aluminum air-cooled 
cylinder operating at maximum load. 



R.p.m 

B.h.p 

Temperature, degrees Ceir 
tigrade 



800 
10.2 

100 



1,000 

12.8 

103 



2,200 1,400 1,600 



15.4 
124 



18 
123 



19.7 
136 



1,800 
20.6 

138 



The compression ratio is more important than engine speed 
in determining the wall temperature. There is usually a definite 
compression ratio giving minimum wall temperature; variations 
on either side increase that temperature. The following table 
gives tests of a 100 by 140 mm. aluminum air-cooled cylinder 
with varying compression ratio. The brake mean effective 
pressure and fuel consumption of this engine are noteworthy. 



Compression ratio 

Brake mean effective pressure, lb. per 

sq. in I 116. 1 

Fuel, pounds per brake horse power per 

hour 

Mean barrel temperature, J Top 

degrees Centigrade \ Bottom 



4.6 


5.0 


5.4 


5.8 


6.2 


116.1 


119.3 


122.0 


125.0 


129.0 


0.530 


0.507 


0.490 


0.475 


0.480 


180 


170 


157 


154 


183 


105 


95 


89 


85 


110 



6.4 

123.0 

0.520 
212 
135 



348 



THE AIRPLANE ENGINE 



The increased temperatures in the last two columns result from 
preignitions which were occasional with compression of 6.2 and 
frequent with 6.4. 

Increase in cylinder diameter diminishes slightly the heat loss 
to the walls per brake horse power; the experimental evidence 
suggests a decrease of about 3.5 per cent for 10 per cent increase 
in cylinder diameter. As the ratio of cooling surface to b.h.p. 
varies inversely as the diameter in similar cylinders, the ratio of 
cooling area to heat given to the walls decreases as the diameter 
increases. The temperature difference between an air-cooled 
cylinder and the cooling air in a given wind may be taken as 
inversely proportional to D - 6 , where D is the cylinder diameter. 

The air -fuel ratio has considerable influence on the heat 
transmitted to the cylinder walls. The cylinder is hottest 
with the weakest mixture capable of sustaining maximum load; 
or, approximately, with an air-fuel ratio of 13.5. Further 
weakening of the mixture makes a cooler cylinder on account of 
the reduction in brake horse power and in the heat loss per 
b.h.p. At the same time it gives a hotter exhaust valve. The 
last point is brought out in the following table giving test 
results for a 100 by 140 mm. air-cooled aluminum cylinder. 



Air-fuel ratio 

Brake m.e.p., lb. per sq. in. . . . 

Fuel, pounds per brake horse 
power per hour 

Exhaust valve temperature, de- 
grees Centigrade 



11.1 


11.9 


13.8 


15.2 


122 


122 


119 


116 


0.622 


0.589 


0.515 


0.480 


706 


717 


747 


752 



15.7 
114 

0.470 

747 



An increase in mixture strength beyond that necessary for 
maximum power reduces the temperature of the cylinder sur- 
faces, as shown below for an engine operating at full throttle and 
constant speed. 



Air-fuel ratio 


10.5 
200 


11.4 
215 


12.9 
237 


13.5 
229 


15.4 


Cylinder head temperature, degrees 
Centigrade 


215 







Tests show that the maximum temperature of the head of an 
air-cooled cylinder must not exceed 270°C. for satisfactory work- 
ing. If the temperature exceeds 280° there is usually trouble 



THE COOLING SYSTEM 349 

from preignition. Higher working temperatures are permissible 
with larger cylinders. If the temperature is kept at 200° to 
220°C, the economy and capacity obtained are quite as good as 
for water-cooled cylinders of similar design and size. 

The temperature of the exhaust valve at its hottest point 
should not exceed 720°C; with valves not exceeding 1.5 in. in 
diameter it is possible to reduce this temperature to 650°C. 

In a well designed aluminum cylinder of the overhead-valve type, 
operating in a 60-mile-per-hour wind, a provision of 0.28 to 0.35 
sq. ft. of cooling surface per brake horse power is sufficient to give 
satisfactory operation, the larger area applying to cylinders of 
about 4-in. bore, and the smaller to cylinders of about 6-in. bore. 
For steel or cast-iron cylinders with overhead valves this area 
must be increased about 50 per cent and for L-head cast-iron 
cylinders by 100 per cent. 

At reduced wind speeds the cylinder temperature increases; 
the mean temperature difference between the fins and the air 
varies inversely as I 70,4 . Thus in a given series of tests a reduc- 
tion of wind speed from 80 to 40 miles per hour increased the 
cylinder temperature from 229° to 296°C. There are practical 
difficulties in the way of providing sufficient cooling surface for 
operation at full throttle below certain limiting wind speeds. 
The wind speeds can be less with smaller cylinders. The mini- 
mum air velocity for good performance of air-cooled cylinders 
of good design and material under full throttle is given below. 
At lower air speeds partial throttle only should be used. 

Diameter, inches 2 3 4 6 8 

Minimum air velocity, miles per hour. 30 40 50 70 90 

Cylinder Materials. — With cylinders of normal design the 
middle portion of the head is the hottest point. Free air flow 
to this point is impeded by the valve ports and gears so that it is 
almost impossible to provide adequate cooling surface there. 
The heat has to travel outward and is dissipated mainly from 
the cooling surface surrounding the combustion head. It 
is therefore important to use a material of maximum thermal 
conductivity. The three practical materials for cylinder con- 
struction, steel, cast-iron and aluminum alloy, have conductivities 
(in C.G.S. units) of 0.12, 0.10 and 0.38 respectively. Aluminum 
is consequently the most desirable material. The alloys most 
suitable for cylinders are copper-aluminum alloys with about 



350 



THE AIRPLANE ENGINE 



90 per cent of aluminum. The high-zinc alloys are unsuitable 
because their tensile strength is low at 200°C. All the alloys 
show rapid decrease of strength as the temperature increases 
beyond 250°C. The following table gives data on this point. 



Composition per cent 


Tensile strength, 
lb. per sq. in. 


Cu 


Sn 


Mg 


Al 


Ni 


Mn 


At 250°C. At 350°C. 



7.0 


1.0 




92.0 






12.0 






88.0 






14.0 






85.0 




1 


4.0 




1.5 


92.5 


2 




9.0 






89.0 


2 





12,300 
23,500 
21,300 
24,600 
19,000 



6,700 
13,400 
14,500 
11,200 
10,100 



Water Cooling. — By water-cooling the cylinder and the 
exhaust ports it is possible to run with higher speeds and com- 
pression ratios than are practicable with air-cooled cylinders. 
The possible increase in speed and ratio of compression are rel- 
atively unimportant when compared with the performance of 
the best recent constructions in air-cooled cylinders; they are 
considerable as compared with the average air-cooled cylinder. 

With water cooling it is possible to maintain almost any 
desired cylinder temperature. If the temperature is low the 
volumetric efficiency and the capacity of the engine will be 
improved (see p. 37) but the engine friction increases and its 
efficiency falls off. The temperature of the jacket water after 
leaving the radiator must be below the boiling point of water at 
the pressure existing on the suction side of the pump, otherwise 
the pump will not function well but will suck in water vapor. 
As fuel economy is ordinarily more important than capacity, 
the jacket water is usually kept at as high a temperature as 
the boiling point will permit. The mean jacket temperature 
at the ground is usually 160 to 180°F. 

Water is not the ideal cooling agent. A less volatile fluid would 
permit higher cylinder temperature; higher efficiencies might be 
obtained without running into such temperatures as would cause 
preignition. The same result might be obtained by operating a 
closed water-cooling system under pressures greater than 
atmospheric, but this would necessitate heavier material for the 



THE COOLING SYSTEM 



351 



radiator core and consequent increase in weight and decrease in 
airplane efficiency. The heat which is removed by the jacket 
water is practically equal to the b.h.p. F or is 42.4 B.t.u. per brake 
horse power per minute. In a closely cowled engine this same 
amount of heat would have to be removed from the radiator. 
With the usual cowling there is considerablefremoval of heat 
by the air stream from the engine and water-jacket surfaces, so 
that only about 31 B.t.u. per brake horse power per minute has 
to be removed from the radiator: with an uncowled engine this 
quality falls to 23 or 25 B.t.u. 

The principal parts of a water-cooling system are the jackets, 
the pump and the radiator. The last of these will be considered 
first. 

Radiators. — Airplane radiators have developed from auto- 
mobile practice but certain types of automobile radiators are 
entirely unsuited to air- 
plane practice. The suc- 
cessful commercial types 
have cores made of thin 
brass, or copper ribbons 
or tubes from 0.004 to 
0.006 in. thick. Common 
types are shown in Fig. 
269, which illustrates: a 
and 6, rectangular air pas- 
sages; c, rhombic passages; 
and d, circular passages with 
hexagonal ends. Other com- 
mon types have hexagonal 
or elliptical air passages 
The water passages are nar- 
row, varying from 0.03 to 
0.08 in. The air tubes are commonly not more than }i in. 
in maximum cross-section dimension and are from 3 to 5 in. 
long (depth of core). The metal sheets are stamped or rolled 
to the desired form with the front and rear ends of the pair of 
sheets forming each water passage in contact with one another. 
These ends are soldered by dipping them into a shallow pool of 
molten solder. Great care must be exercised to keep down the 
weight of solder as much as possible; it often amounts to 25 per 
cent of the total weight of the radiator core. The top and 




"Water 

(d) 

Fig. 269. — Types of radiator core. 



352 THE AIRPLANE ENGINE 

bottom ends of the water passages are inserted through slots in 
the top and bottom headers respectively; the two sheets of each 
water passage are spread apart and soldered to the header. 
In the case of type d, Fig. 269, the ends of the circular tubes are 
expanded into hexagonal forms which are soldered together; the 
expansion is made enough to give the desired width of water 
passage between the tubes. Type b differs from type a not 
only in the method of assembly but may also be made of a cor- 
rugated surface which is intended to give greater strength and 
larger radiating surface. The types a, b, c and d in Fig. 269 have 
water in contact with all the radiating surface and are said to 
have only " direct" radiating surface. Many automobile 
radiators have extensions of this direct radiating surface in the 
form of fins on flat or circular tubes (e, Fig. 269), metal spirals, 
and so forth. Such ''indirect" radiating surface is found to 
have too high a ratio of head resistance to heat-removing capacity 
to be satisfactory for airplane use. 

The dimensions or external shape of a radiator can be adapted 
to suit its location and desired performance. The location 
may be such that air may pass through or around it without 
obstruction, in which case it is said to be in an "unobstructed" 
position. On the other hand, the radiator may be located in the 
nose of the fuselage, or in the plane of the wing, in which case 
the air flow is materially affected by other parts of the plane 
and the radiator is said to be "obstructed." The performance of 
such a radiator will depend not only on the size and type of the 
core but on its position or surroundings. Examples of typical 
unobstructed locations are shown in Fig. 270 (at the sides of 
fuselage) and in Fig. 271 (over the engine); common obstructed 
positions are in the nose of the fuselage, and in the wing. 

A comprehensive study of the properties of various types and 
dimensions of radiator cores has been made at the Bureau of 
Standards and published in the Fifth Annual Report of the 
National Advisory Committee on Aeronautics. The following 
discussion is mainly from that source. 

Two quantities are of importance in determining the heat 
transfer of a core. They are the temperature difference between 
the entering air and the mean water temperature; and the 
mass flow of air. The temperature difference should ordinarily 
be taken as the difference between the mean summer air tempera- 
ture and the mean water temperature. The mass flow of air, 



THE COOLING SYSTEM 



353 



Outlet P/'pe^ 




Fig. 270. — Side radiators. 





jtxl 



Thermometer*'---'' 



Fig. 271. — Overhead radiators. 



26 



354 THE AIRPLANE ENGINE 

M, is the weight of air flowing per second per square foot of 
frontal area of the core. Its amount (at constant air density) 
is found to be proportional to the free air speed or the velocity 
with which the core moves through the air when the core is 
unobstructed; the mass flow is always less for obstructed positions 
than for unobstructed. 

The energy dissipated or heat transfer is expressed in horse 
power per square foot of frontal area, and, for purposes of com- 
parison of the properties of various cores, a temperature difference 
of 100°F. is assumed between the air entering the radiator and the 
mean water temperature; the heat transfer is proportional to this 
temperature difference. One horse power is equivalent to 42.54 
B.t.u. per minute. 

The head resistance of the core is the force required to push it 
through the air and is expressed in pounds per square foot of 
frontal area. This head resistance is found to vary approxi- 
mately as the square of the free air speed; in most cases the 
exponent is slightly less than 2. If R is the head resistance, 
and V the free air speed in miles per hour, then 

R = cV 2 

and c is called the head resistance constant. 

The horse power absorbed by a radiator is the engine power 
required to overcome the head resistance and support the weight 
of the radiator. The work done in supporting the weight can be 
calculated if the lift-drag ratio of the plane as a whole is known. 
An average value of 5.4 may be assumed for this ratio. If W 
is the weight of the core and contained water in pounds per 
square foot of frontal area, the propeller thrust required to sup- 
port the weight is W/5A. The horse power absorbed is 

H.P. = IB + ")■ V X 5 ' 280 



P = (R + - 

\ T 5.4/ 60 X 33,000 



It should be noted that this method of calculation neglects the 
effect on the lift-drag ratio of the addition of the radiator. 
The lift-drag ratio varies between different planes and varies 
even more widely between climbing and level flight. The selec- 
tion of a core for a given plane cannot be made satisfactorily 
without a knowledge of the relative importance of climbing 
speed and top speed. A lift-drag value of 5.4 is a good average 



THE COOLING SYSTEM 355 

and gives about equal value to climbing and level speed. If the 
rate of climb is of prime importance the value may be as low as 
3, while if speed on the level is the most important, a value as 
high as 10 may be used. 

A small additional power charge against the radiator is that 
required to overcome the resistance to water circulation in the 
radiator. It is usually so small as to be negligible. 

The definition just given of horse power absorbed applies only 
to the case of an unobstructed radiator. If the addition of a 
radiator necessitates alterations in structure (such as the sub- 
stitution of a flat nose for a stream-line fuselage or the enlarge- 
ment of the fuselage to accommodate the radiator required) 
the consequent increase in resistance of the structure should be 
charged to the radiator. 

A comparison of the performance of various cores can be 
obtained when the heat transfer per unit of power absorbed 
is known. This quantity is called the Figure of Merit and is 
a pure number. The comparison must be for the same tempera- 
ture difference and free air speed. It applies only to unob- 
structed radiators. 

The general conclusions derived from the tests at the Bureau 
of Standards are as follows: 

Heat transfer is a function of mass flow of air and is inde- 
pendent of the air density. 

Heat transfer is roughly proportional to mass flow for a core 
having only direct cooling surface. When there is a considerable 
amount of indirect cooling surface the heat transfer increases less 
rapidly than mass flow at high air speeds. 

Heat transfer is proportional to the temperature difference. 

Heat transfer is not greatly affected by the rate of water flow pro- 
vided the rate is above 2 gal. per minute per inch of core depth per 
foot width of core. It should be noted, however, that this is true 
only when the mean water temperature is regarded as constant. 

Heat transfer from direct cooling surface is not appreciably 
affected by the composition of the metal. When fins and other 
indirect cooling surface are used the thermal conductivity of the 
metal is important. 

Heat transfer is somewhat increased, but at the expense of a 
large increase in head resistance, by spirals or other forms of 
passages which increase the turbulence of the air stream. Heat 
transfer is greater for smooth than for rough tube walls, for, if 



356 THE AIRPLANE ENGINE 

the surface is rough, it will be covered with a layer of more or less 
stagnant fluid. 

Head resistance for any particular core varies approximately 
as the square of the free air speed. 

The head resistance of a core appears to be closely related to 
its mass flow so that, in general, anything which tends to cut 
down the flow of air through the core will cause a considerable 
increase in head resistance. 

Head resistance varies directly as the air density for a given 
free air speed, and inversely as the density for a given mass flow. 

Head resistance is considerably increased by projections, 
indentations, or holes in the air tube walls. 

Head resistance per square foot is not appreciably affected by 
the size of the core within the limits used, viz., 8 by 8 in. to 16 by 
16 in. and 12 by 24 in. 

Special conclusions with reference to types of cores are as 
follows: 

For a high figure of merit the core should have smooth, straight 
air passages, easy entrances and exits for the air and a large per- 
centage of free area. Under these conditions the figure of merit 
increases as the depth increases up to at least 20 times the di- 
ameter of the air tubes, which is as far as experiment has gone. 
Even greater depths may be of advantage. 

By far the most satisfactory radiator for use in unobstructed 
positions seems to be one of thin flat plates with water space not 
over }^6 in. wide and spaced % in. on centers. The plate should 
be at least 12 in. deep. As the figure of merit changes but 
slightly with increase of depth beyond 12 in. the depth may be 
made 20 in. or more if it is desirable to reduce frontal area. The 
chief defect of the type is mechanical weakness. Of the commer- 
cial radiators tested, those have given highest figure of merit at 
high air speeds which have only direct cooling surface in the 
form of tubes about }4 in. square and about 5 in. deep. The 
figure of merit of this type at 120 miles per hour free air speed 
varies from about 8 to 8.4, whereas flat plates 9% in. deep and 
3^ in. on centers have a value of 10.7. The energy dissipated per 
square foot of frontal area is less in the above flat plate radiator 
than in the best square tube radiators so that a larger frontal 
area will be required with flat plates but the power absorbed 
will be less. 

The British Air Ministry has adopted as standard a circular 



THE COOLING SYSTEM 357 

tube 10 mm. in diameter expanded at the ends to a hexagonal 
section 11 mm. across the flats (Fig. 269 d). The standard length 
of the tubes is 120 mm. The material is 70 — 30 brass with wall 
thickness of 0.005 in. 

The actual power absorbed by the radiator in being lifted and 
pushed through the air (see Table 18) varies from about 3 h.p. 
to 6 h.p. per square foot of frontal area at 100 miles per hour. 
This amounts to from 5 to 20 per cent of the total engine power. 
A small gain in radiator performance may have an appreciable 
effect at high speed. 

The selection of a radiator core for an obstructed position is 
more difficult. An obstructed position involves a large absorp- 
tion of power. The resistance of a fuselage fitted with a nose 
radiator is two or three times the resistance of the same fuselage 
with a stream-line nose. The increase in resistance due to the 
substitution of a radiator for a stream-line nose is greater than 
the increase that would be caused by using a radiator of the 
same core construction and the same cooling capacity in an 
unobstructed position. 

At any given free-air speed the total resistance of a fuselage 
with a flat nose radiator is increased by increasing the air flow 
through the radiator, either by opening exit vents for the air or 
by decreasing the resistance of the radiator to the passage of air. 
This indicates that a nose radiator should be of compact con- 
struction with high heat transfer, for low air flows through the 
core, requiring a core of high resistance. This fact is of special 
importance since the space available for a nose radiator is so 
limited that the highest possible mass flows are used in practice. 
A nose radiator with air exit vents equal in area to the free air 
passage through the radiator is found to cut down the heat 
transfer about 35 per cent as compared with the same radiator 
in an unobstructed position. Indirect cooling surface may be of 
advantage if it is made of copper, crimped from the water tube 
walls and well soldered to them at all possible places. Several 
types of core show good heat transfer at low speeds, but here 
again the square tube, with direct radiating surface only, gives 
best result of all commercial types, and flat plates spaced J4 in. 
on centers show excellent performance. The fin and tube type 
with its small amount of direct surface has no use in airplanes 
except possibly in a wing position where a high head resistance 
is no disadvantage. 



358 



THE AIRPLANE ENGINE 



The properties of cores selected as typical of common con- 
struction are given in Tables 17 and 18. The constructions vary 
from the flat tubes (E-Q and E-S) with 100 per cent direct cooling 
surface and a minimum of head resistance to finned circular 
tubes (^-5) with only 12.3 per cent of direct surface and three 
times as much head resistance per square foot of frontal area as 
the best flat tubes. The dimensions of the core are given in Table 
17; the performance at various wind speeds in unobstructed posi- 
tions in Table 18. The "figure of merit" necessarily diminishes 



^ ^ 

1 

01 — 111 — I I 1 , 1 — _ 



0.020 0.040 0.060 0.080 0.100 

Mass Flow Fac+or (k) 

Fig. 272. — Head resistance constant and mass flow factor. 



with increase of wind speed; the order of merit of the different 
cores is different at different speeds. At 120 miles per hour the 
flat tubes E-S and E-Q are seen to be best and the finned circular 
tubes F-5 the poorest. 

Table 19 gives the constants in the empirical equations for 
"Head Resistance," "Mass Flow" and "Energy Dissipated" for 
the cores listed in Table 17; these quantities are obtained from 
the data of Table 18. The Head Resistance Constant and the 
Mass Flow Constant appear to be connected by a simple relation; 
plotting these quantities for all the cores tested gives the curve of 
Fig. 272. 



THE COOLING SYSTEM 



359 



OftOiON 



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360 



THE AIRPLANE ENGINE 



Table 18. — Radiator Performance in Terms of Free Air Speed (for 
Unobstructed Positions Only) 
Grade A represents very good performance; grade E, very poor 



Radiator 



Speed, 
miles 
per 
hour 



Air flow, 
lb. per 
sq. ft. 

per sec. 



Energy in h.p. 
dissipated per 
square foot of 



Front Surface 



Head 
resistance, 
lb. per 
sq. ft. 
frontal 
area 



H.p. 

absorbed 

per sq. 

ft. of 

frontal 

area 



Figure 
of merit 



A-7. 




30 

60 

90 

120 



30 

60 

90 

120 



30 

60 

90 

120 



30 



90 
120 



30 

60 

90 

120 



30 

60 

90 

120 



30 

60 

90 

120 

30 

60 

90 

120 



30 

60 

90 

120 



Grade. 



2.20 
4.40 
6.60 

8.80 



2.29 
4.58 
6.87 
9.16 



2.12 
4.24 
6.36 

8.48 



2.40 
4.80 
7.20 
9.60 



2.41 
4.82 
7.22 
9.63 



2.12 
4.24 
6.36 

8.48 



2.74 

5.48 

8.23 

10.97 

1.82 
3.64 
5.46 

7.28 



1.88 
3.75 
5.62 
7.50 



27.2 
45.9 
61.7 
76.5 



20.3 
37.2 
52.9 
68.0 

B 
20.2 
31.0 

39.8 
47.9 

D 

14.8 
26.4 

37.8 
48.5 

D 
13.8 
24.7 
32.3 

38.7 

D 

29.3 
51.1 
71.3 
90.5 

A 
17.2 
29.8 
41.0 
51.7 

C 
13.8 
21.7 
27.8 
33.0 

E 

23.1 
39.5 
53.6 
67.0 



0.43 
0.73 
0.99 
1.22 



0.37 
0.68 
0.97 
1.25 



0.36 
0.55 
0.71 
0.85 



0.46 
0.83 
1.18 
1.52 



0.58 
1.04 
1.32 
1.63 



0.37 
0.65 
0.91 
1.15 



0.44 
0.76 
1.05 
1.32 

0.40 
0.63 
0.81 
0.96 



0.52 
0.88 
1.20 
1.50 



1.72 
6.70 
14.90 
26.3 

D 

1.40 
5.61 
12.10 
21.3 

C 

1.72 
6.88 
15.48 
27.5 

D 

1.37 
5.47 
12.30 
21.9 

C 

1.25 
4.75 
10.48 
18.4 

B 

1.57 
6.27 
14.10 
25.1 

D 

0.78 
3.12 
7.03 
12.5 
A 

2.52 
9.65 
21.6 
38.3 

E 

2.40 
8.65 
19.1 
33.4 

E 



.50 
1.79 
4.66 
9.87 

D 
0.44 

1.54 
3.88 
8.11 

C 

0.41 

1.64 

4.52 

9.89 

D 

0.30 
1.26 
3.53 

7.79 

C 

0.26 

1.08 

2.99 

6.53 

B 

0.51 

1.78 

4.54 

9.59 

D 

0.27 
0.92 
2.31 
4.84 
A 

0.33 
1.81 
5.57 
12.8 

E 

0.40 
1.80 
5.21 
11.5 

E 



54.6 
25.7 
13.2 

7.8 

B 
46.6 
24.1 
13.6 

8.4 

B 

49.8 

18.9 

8.8 

4.8 

D 

48.7 

20.9 

10.7 

6.2 

C 
53.5 
23.0 
10.8 

5.9 

C 

57.2 

28.8 

15.7 

9.5 

B 
63.3 
32.4 
17.8 
10.7 

A 
41.3 
12.0 

5.0 

2.6 

E 
58.0 
22.0 
10.3 

5.8 



THE COOLING SYSTEM 



361 



A plotting of the data in Table 18 for core E-S is given in 
Fig. 273. 

Size of Radiator. — If a type of core has been selected and its 
properties are known and plotted as in Fig. 273, the necessary 
size of an unobstructed core is directly obtainable. The dimen- 
sions must be calculated for the most unfavorable condition, 
which, ordinarily, will be maximum climbing speed near the 
ground with summer temperatures. The mean water temper- 
ature will be fixed by the maximum temperature allowable in 
the jackets and by the quantity of water circulated. The 





D 




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50 


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Hr. 


po 






% 




£ 


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^^ />7 Dee/ 
39.2Sq.Fr.i 


1 ' 












28 








Surface 














100 


/" Direct Surface 








24 










3.2 lb Fmphf 








Merit, r 

~>. Absor 

o 










S.D lb. Wafer 








20 
























C fe 40 


















































i. . 30 
8-i: 
























12 














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10 



2 4 6 8 

Mass Flow of Air, Lb. perSq.Ft.perSec. 
Fig. 273. — Properties of a flat tube radiator core. 

maximum allowable water temperature is ordinarily 20 to 30°F. 
below the boiling point and varies with the altitude of the plane ; 
its value will be determined by (1) its influence on the volumetric 
efficiency of the engine (see p. 37) and (2) the water resistance 
of the radiator (see p. 364). If the water system is closed, 
with no vent to the atmosphere, the last-mentioned factor 
disappears. The heat to be dissipated should be determined if 
possible but may be assumed equal to the brake work of the 
engine if more exact knowledge is not obtainable. The effect 
of propeller slip should be estimated and allowed for. Allowance 
should also be made for the cooling effect of the radiator headers 
and of the exposure of the engine to the wind. 



362 



THE AIRPLANE ENGINE 



Occasionally, instead of designing for maximum climb some 
other condition may impose maximum service on the radiator 
as, for example, in flying boats and seaplanes intended for 
training, where much taxi-ing is done at low plane speed and 
maximum engine power. If the radiator is in the nose of the 
fuselage some assumption must be made as to the relation of mass 
flow to the speed of the plane. The mass flow will usually vary 
from 0.04 to 0.07 time the speed of the plane (in miles per hour), 
depending on the type of radiator, the amount of cooling and the 
masking effect of the propeller. The power absorbed is seldom 
calculable because of the uncertain effect of the radiator on the 
resistance of the fuselage. 

Table 19. — Constants in the Equations R = cV 2 ; M = kV; and Q = Gm n 

R = Head resistance in pounds per square foot. 

V = Free-air speed in miles per hour. 
M = Mass flow of air in pounds per second per square foot. 

Q = Energy dissipated in horsepower per square foot per 100°F. tem- 
perature difference. 

m = "mass flow constant," which is the ratio of the mass of air passing 
through 1 square foot of radiator to the mass of air passing through 
1 square foot of free area in front of the radiator. 



Radiator 


c X 10 3 


k X 10 2 


m 


G 


n 


A-7 


1.86 

1.56 

1.91 

1.52 

1.32 

1.74 

0.867 

2.68 

2.40 


7.34 
7.63 
7.07 
8.00 
8.03 
7.07 
9.13 
6.07 
6.25 


0.667 
0.694 
0.643 
0.727 
0.730 
0.643 
0.830 
0.552 
0.568 


15.1 

10.3 

13.1 

7.1 

8.1 

16.1 

7.6 

10.0 

14.8 


0.75 


A -23 


0.85 


B-S 

C-4 


0.60 ■ 
0.85 


D-l 

E-Q 


0.70 
0.80 


E-S 


0.80 


F-5 


0.60 


G-3. 


0.75 







The mass flow for a wing radiator depends on the angle of 
incidence but is probably not over 0.01 time the plane speed 
even at the best climbing angle. 

The relative efficiencies of radiators in various positions are 
given by Liptrot 1 as follows: 

1 Aeronautics, Apr. 29, 1920. 



THE COOLING SYSTEM 363 



Position of radiator 



Relative 
efficiency 



Unobstructed 

Underslung, side or overhead, but close to fuselage 

Twin nose radiator 

Nose radiator with core entirely above or below propeller 

shaft 

Nose radiator with propeller shaft in center 

Behind engine 



1.000 
0.973 
0.716 

0.656 
0.585 
0.423 



The use of a small projecting lip or stream line entrance around 
the core may reduce the necessary core size slightly but at the cost 
of a considerable increase of head resistance. 

Rate of Water Flow. — One gallon (231 cu. in.) of water at 

60.12 X 231 
200°F. weighs — ~Yt>r = 8 lb. approximately. With a 

temperature difference of 10°F., 1 gal. of water per minute 

80 
will give up 80 B.t.u. or t^TTk = 2 h.p. approximately. With a 

temperature difference of 5°F. the flow of water in gallons per 
minute should equal the engine horse power. 

The entering temperature of the water is fixed by the necessity 
of keeping at a certain point below boiling. With fixed entering 
temperature, if the amount of water circulated is increased the 
mean temperature of the water is raised and consequently the 
temperature difference between air and water is increased. The 
influence of water velocity on the heat transfer is found by 
experiment to be very small so long as the velocity is above 2 gal. 
per minute per foot width per inch depth of core, which is much 
below usual rates. With a circulation of 34 gal. per minute per 
horse power the temperature fall of the water is 20°F. ; increasing 
this to y% gal. reduces the temperature fall of the water to 10°F. 
and increases the temperature difference between air and water 
by 5°F. With an infinite amount of water circulated this tem- 
perature difference could be increased only another 5°F. The 
increase in pump work with increased water flow makes it 
undesirable to circulate more than about J-£ gal. per minute per 
horse power, and with radiators that are relatively long and 
narrow a flow of 34 gal. per minute per horse power should be 
used. 



364 



THE AIRPLANE ENGINE 



The pressures required to maintain water flow through the 
cores of radiators vary greatly with the dimensions and type of 
construction. Those types having the widest and straightest 
water spaces offer least resistance whereas those with many right 
angle bends will offer much resistance. The range for 12 com- 
mercial radiators tested at the Bureau of Standards, all of them 
8 in. square in frontal section and of depths varying from 2% to 
4 in. with a total water flow of 20 gal. per minute, was from 0.27 
to 10.2 ft. of water pressure drop. These pressure drops may be 
assumed to vary directly as the height of the core, but the rate of 
change with change of water velocity follows an exponential law 
in all cases, though with a widely varying exponent in the 
different types. The resistance seems to depend largely on the 
care used in manufacture and on the form of the water tube 
entrances and exits. It would seem well to include a test for 
pressure necessary to produce water flow in acceptance specifi- 
cations for complete radiators. 

The water enters the top header of the radiator, at which place 
atmospheric pressure is usually maintained through the overflow 
pipe. The suction pressure at the pump cannot be less than the 

vapor pressure of the 

Tank on Top Plane , ., , 

water leaving the rad- 
iator if the pump and 
radiator are at the same 
level. If the water leaves 
the radiator at 190°F. 
the corresponding vapor 
pressure is 9.2 lb. or 
about 5 lb. below at- 
mospheric pressure. The 
maximum pressure avail- 
able for overcoming the 
resistance of the radiator in this case will be 5 lb. per square inch 
or 11.5 ft. of water. With a reserve tank in the upper plane, as 
in Fig. 274, the head available in overcoming radiator friction 
is increased by the height of the tank above the suction. If 
the resistance of a proposed radiator is in excess of the available 
pressure, its height must be decreased and its width correspond- 
ingly increased in order to give the necessary radiating surface. 

Occasionally radiation or expansion tanks instead of being 
vented to the atmosphere are provided with safety valves opening 




Carburetors 



Fig. 274.- 



Radiahrs on each side 
of Machine 
-Cooling system of Benz engine. 



THE COOLING SYSTEM 



365 



at 2 or 3 lb. per square inch. This diminishes the loss of water 
from evaporation and may permit a higher water temperature. 
Effect of Altitude on Radiator Performance. — The investiga- 
tions at the Bureau of Standards have yielded the following 
general conclusions: 

The effect of the lower air temperature is to increase the heat 
transfer in proportion to the increase in the mean temperature 
difference between the entering air and the water. The decrease 
in air density reduces the mass flow of air and decreases the heat 
transfer at any given plane speed in proportion to the air density. 
Head resistance is proportional to air density and is therefore 
reduced with increased altitude. The combined effect of temper- 
ature and density changes is to decrease the heat transfer but 
not as rapidly as the engine power diminishes; consequently 
the cooling capacity of the radiator becomes excessive at high 
altitudes and may be more than double the required capacity. 
As the head resistance 
falls off more rapidly than 
the heat transfer the 
" figure of merit" of the 
radiator increases with 
altitude. 

From the above con- 
clusions the performance 
of a radiator at any alti- 
tude can be calculated 
when its ground perform- 
ance is known. For ex- 
ample, take the flat plate 
core (E-8) for which 
ground data are given in 
Tables 17 and 18. It is 
desired to calculate its performance in summer at 10,000 ft. 
altitude and 120 miles per hour. The ground data are: 

Mass flow of air at 120 miles per hour = 10.97 lb. per square 
foot per second. 

Head resistance at 120 miles per hour = 12.5 lb. per square 
foot. 

Weight of core and contained water = 14.15 lb. per square 
foot. 

The mean temperature of the water in the radiator may be 



50 



u-. '40 

gno 



g-20 



-40 



-60 























^ 


^ 














-%' 


•%• 














^*vf- 






















\ 


> 


X 



































Fig. 275. 



I- 8 12 1G 20 24 28 

Altitude inThousands of Feet 
-Variation of air temperatures with 
altitude. 



366 



THE AIRPLANE ENGINE 



assumed to be 30°F. below the boiling point. The pressure at 
10,000 ft. is 10.2 lb. per square inch (see p. 389) and the cor- 
responding boiling point is 194. 2°F. The summer mean temper- 
ature at 10,000 ft. may be taken as 45°F. (Fig. 275). The 

mean temperature differ- 
ence at 10,000 ft. will be 
194.2 - 30 - 45 = 
119.2°F. The air den- 
sity at the same ele- 
vation is 0.0545 lb. per 
cubic foot (see Fig. 276). 
The mass flow at 10,000 

ft. = 10.97 X £gg| = 

7.981b. per square foot 
per second. The energy 
dissipated at mass flow 
of 7.98 lb. is 40 h.p. per 

square foot per 100° F. temperature difference (see Table 18); 

with the increased temperature difference the energy dissipated 
119.2 




Fig. 276.- 



5 12 16 20 24 28 
Altitude in Thousand&of Fee+ 

-Variation of air densities with 
altitude. 



becomes 40 X 



100 



= 47>.7 h.p. per square foot. The head 

0.0545 
12.5 X qqjzq = 9.09 lb. per square 



resistance (see Table 18) 

foot. 

The degree of masking required at altitudes may be readily 
calculated if the engine h.p. is assumed proportional to the air 
density. If the radiator is 
just adequate in level flight 
at a given speed at the 
ground, it will be capable of 
more cooling than is required 
of it in level flight at the same 
speed at higher altitudes. It 
is therefore possible to mask 
an increasing fraction (and 
cut down thereby the mass 
flow) as altitude increases. 
The curve of Fig. 277 shows how much masking is possible for 
the flat plate radiator E-8 at 120 miles per hour, but the curve 
is practically the same for other cores and speeds. 



100 



80 



h 40 



20 































E-8, 














6-3 





































4 8 12 16 20 24 
Altitude in Thousands of Feef 



28 



Fig. 277. — Radiator masking at altitudes. 



THE COOLING SYSTEM 



367 



It should be remembered that climbing should be considered 
as well as level flight in any discussion of radiators and of mask- 
ing. The speed for maximum climb may be only one-half that 
of level flight at certain altitudes, and the cooling must be 
adequate for the climbing condition. This consideration alone 
would require a masking of 50 per cent for such planes in level 
flight. If the relation between maximum climbing speed and 
level speed is known, and also the change in engine revolutions 
and power, the mass flow of air can be determined under both 
conditions and the desirable degree of masking can be found. 



Valve 



Pump 





Pump 



Jo connect systems 
in series, furn A" 
ih rough 90° 



Engine ' 



Radiator 



Fig. 278. — Radiator interconnections for dual engines on lighter-than-air 

machines. 



The twin-engined dirigible offers a special case of importance. 
Such a craft may operate for long periods with one engine only, 
which therefore operates at low speed but full power. If the 
radiator is designed for maximum speed each radiator will be 
too small for its engine at this reduced speed. To obviate the 
use of a larger radiator the installation may be arranged as in 
Fig. 278 in case the water pumps are of such construction as to 
permit the water to pass through when they are idle. Turning 
the valve A through 90 deg. will circulate the water through 
both radiators and through the jackets of both engines, and will 
thereby prevent the idle engine from freezing up and will give 
more than adequate radiating surface. Some provision for 
masking the radiator is especially desirable in this case. 



368 THE AIRPLANE ENGINE 

Masking can be partially accomplished by varying the water 
flow, as by by-passing some of the water from radiator inlet to 
outlet. The effect of reducing the quantity of water is to reduce 
the mean temperature of the water and thereby to reduce the 
mean temperature difference between air and water. The 
possible range of control by this means is small. Shutters across 
the radiator front answer the purpose more fully, although 
they add to the head resistance. They may be operated by the 
pilot, or as in some German planes, may be under the automatic 
control of an electrical resistance thermometer. Closed shutters 
on a nose radiator decrease the head resistance: on a free air 
radiator, they increase it. A retractable side or bottom radia- 
tor, which may be drawn within the body to decrease the cooling 
effect, is occasionally used. It may be arranged most conven- 
iently as an auxiliary radiator in series with a fixed main radiator 
which has no masking device and is adequate for high-altitude 
evel flight. The auxiliary radiator is retracted as altitude 
is gained. The increased water resistance from two radiators 
in series is objectionable. Yawing is another possibility. 

Effects of Yawing Airplane Radiators. — The air stream does 
not always approach the radiator at right angles to its face. 
The most common causes of this are : 

1. Radiator mounted in the propeller slip stream where the air strikes the 
radiator at angles other than normal to its face. 

2. Radiator mounted in the wing (or other position) where the axes of its 
passages for the air are not parallel to the direction of motion of the plane. 

3. Radiator pivoted about an axis perpendicular to the direction of motion 
of the airplane for the purpose of changing its inclination for the regulation 
of cooling capacity (masking). 

The effects of yawing a radiator through angles from to 
45 deg. are (1) to decrease slightly the mass flow; (2) increase 
the head resistance by as much as 50 per cent in the case of cores 
of low head resistance but much less in the case of high-resistance 
cores; and (3) in some cases, for angles up to 20 or 25 deg. to 
increase slightly the heat transfer. These effects vary largely 
with different types. 

The complete radiator consists not only of the core but of 
top and bottom headers. The top header may serve merely as a 
distributor or it may have sufficient capacity to serve as reserve 
and expansion tank also. The latter practice reduces complica- 
tions and is therefore used on small machines intended for short 



THE COOLING SYSTEM 



369 



flights. For large machines used for long flights, an adequate 
water capacity would entail a large frontal surface and excessive 
head resistance of the header. The desirable reserve capacity 
in British practice is given by the formula: 



Gallons 



h.p. X (endurance in hours) 
1,600 




cccoooooo 
fclooooooooooo 



3— Holes 



J^ 



Baffle Pla+e 
'ns+rumenf ^ mfgfm 

Flanqe---.^ 



Baffle Plate 
Ink __i, Wl 



Uppe, 
Tank 



-Jf Shutter 
Bracket 



Supporting 
Brackets 




Fig. 279. — Details of typical nose radiator. 

The reserve water tank is often located in the upper wing but 
there is danger of freezing unless, as in Fig. 274, the water circula- 
tion is through the tank; the objection to including it in the 
circulation is the increased length of pipe through which the 
water has to be forced. 

The lower header is a collector only and should be as small as 
practicable. Both headers should be stream-lined. The headers 
and their contents will usually add 50 per cent to the weight 
of the core and its contents. Occasionally (as in the Maybach 

24 



370 THE AIRPLANE ENGINE 

plant) the headers are divided into halves by vertical baffles 
on the fore and aft line. Water enters the left-hand side of 
the lower header, passes to the left-hand side of the upper header, 
then over the baffle to the right-hand side and down to its exit at 
the right-hand side of the lower baffle; this arrangement causes 
greatly increased water resistance if the same weight of water is 
circulated; if the weight of water is halved so as to maintain the 
same velocity in the radiator passage, the mean temperature 
difference between air and water will be diminished, necessitating 
the use of a large radiator. 

A complete nose radiator is shown in Fig. 279. Among the 
details to be noted are the filler, inlet and outlet pipes; the 
perforated baffle plate between the inlet and the upper tank; 
the overflow pipe; the upper and lower supporting brackets; and 
the shutter brackets. The filler cap is of hard rubber with a 
safety chain; a better construction, avoiding loss from the snap- 
ping of the chain, is with a hinged cap held closed by a snap wire. 

Pumps. — As previously pointed out (p. 363) the cooling water 
required is not more than 3^ gal. per brake horse power per 
minute. Ordinarily it is Ji gal. per minute or less. The re- 
sistance to the circulation of the water is chiefly in the radiator, 
but is considerable in other parts of the system; its magnitude is 
variable, but may be assumed to be from 4 to 8 lb. per square 
inch in good installations. 

The water horse power of the pump of a 100-h.p. engine using 

Ji gal. (2 lb.) of water per horse power per minute against 8 lb. 

2 X 100 1 

per square inch pressure is 8 X 144 X — ™ — X oo qqq = 

0.116 h.p. If the efficiency of the pump and its drive is 20 per 
cent, the horse power used to drive the pump will be 0.166 -f- 
0.2 = 0.58 h.p., which is a very small fraction of 100 h.p. 
Consequently, the water pump efficiency is comparatively 
unimportant and the type selected should be one of maximum 
simplicity and minimum weight. The single impeller volute 
centrifugal pump meets these conditions best and is univer- 
sally used. 

In a volute pump, water enters axially, is caught by the im- 
peller blades and is given a high velocity of rotation before it is 
discharged into the volute casing, from which it escapes through 
one or more outlets. The number of outlets is usually the same 
as the number of banks of cylinders. In order to keep down the 



THE COOLING SYSTEM 371 

size and weight of the pump the impeller rotates at a speed greater 
than that of the engine; one and one-half engine speed is common. 
If the impeller blades are radial (Fig. 280) the theoretical dis- 
charge pressure in feet of water is given by V 2 /2g where V is the 
tip speed of the impeller. Taking an impeller diameter of 4 in. 

/ 4 2 400\ 2 
and a speed of 2,400 r.p.m. this becomes (7rXjnX "kft/ ~*~ 

2g = 32 ft. = 13.8 lb. per square inch. The water velocity 
leaving the impeller cannot be converted completely into pressure 
head and there are various impeller and casing losses, so that the 



Outlet' 



Inlet - 



Fig. 280. — Water pump of Liberty engine. 

actual discharge pressure will be much less than that calculated 
above; it would probably be less than one-half the theoretical 
value. 

The resistance to be overcome by the pump is entirely fric- 
tional, and varies as the square of the amount of water circulated. 
The amount of water circulated is proportional to the speed of the 
pump. The work done by the pump is proportional to the 
volume of water circulated multiplied by the resistance, or is 
proportional to the cube of the pump speed. 

The pump of the Liberty engine, Fig. 280, has a 2-in. inlet 
and two outlets. It runs at l}^ times engine speed, and has a 
capacity of 86 gal. per minute at 2,000 r.p.m. of the engine. The 
impellers are radial and are partly shrouded. The packing of the 



372 



THE AIRPLANE ENGINE 



impeller shaft against water leakage is kept compressed by a 
coiled spring. The King-Bugatti (410 h.p.) pump (Fig. 281) 
has impellers completely shrouded on one side. As there is only 
one outlet the casing is of complete volute form. The impeller 
is 5% in. diameter with eight vanes, the web being drilled to 
equalize the water pressure. The shaft is packed with graphited 
asbestos rope packing, held under compression by a coiled spring. 
The inlet is 2Ji in. in diameter; the outlet is 2% 6 m - It is 
coupled direct to the engine shaft. 




Fig. 281. — Water pump of Bugatti engine. 



The Austro-Daimler (200 h.p.) pump weighs 7.6 lb., has an 
impeller 4.4. in. in diameter, inlet and outlet diameters 1.42 
in., and a ratio of pump to engine speed of 1.89. It is driven from 
the rear end of the crankshaft by a bevel gear which is integral 
with a sleeve forming an extension shaft (Fig. 282). The pump 
bevel gear floats on the end of the pump spindle, and is fitted 
with a large-diameter thrust ball-race and retaining spring, which, 
being at the bottom end of the spindle, are as far away as possible 
from the impeller. Both the pump spindle bearings are lubricated 
through two holes drilled in the pump body and oil grooves cut 
in the spindle bearings. The impeller is formed with six vanes 



THE COOLING SYSTEM 



373 



and is completely shrouded; it is keyed to the spindle and secured 
by a gun-metal nut and washer. A conically-faced shoulder is 



Inlzt^ 




Fig. 282. — Water pump of Austro-Daimler engine. 



machined on the pump directly beneath the impeller. This 
shoulder beds into the bevelled face of the bronze bearing, form- 



so 

I 40 

to 

^30 

































j\Vf 


.*> 




S*£ 


<< 


L 






W 1 


<£ 


s"2 






Afrg 




6^> 








o^ 










<Q 

































1000 



1200 



1500 



IQQQ 



1400 1600 

Pump R.D.m 
Fig. 283. — Performance curves of Austro-Daimler water pump. 

ing a water-tight joint. The performance curves of this pump 
are given in Fig. 283. 



374 



THE AIRPLANE ENGINE 



The Maybach (300 h.p.) pump (Fig. 284) has an impeller 
4.46 in. in diameter, inlet 2.13 in. in diameter, outlet 1.97 in. in 
diameter, and a ratio of pump to engine speed of 2. The pump 

spindle is driven through a 
dog clutch at its lower end 
by a short vertical spindle 
running in a bronze bush- 
ing; this spindle is driven 
by a bevel gear meshing 
with the main bevel fixed 
on the rear end of the 
crankshaft. The top por- 
tion of the pump spindle 
bearing is cupped to form 
the housing for a thrust 
ball-race, above which is 
fixed the impeller. The 
impeller is a gun-metal 
casting, having six helical 
vanes. The lower half of 

-Water pump of Maybach engine. the pump body ig an 

aluminum casting, to the inlet passage of which the diagonal 
water pipe from the radiator is coupled by a rubber con- 
nection. The top half of the water pump body, which is a gun- 
metal casting, is formed with six helical passages leading in 




Fig. 284. 



£|24 
pVfc20 



^30 



















. 
















J^Sc 


/W 


^j&& 




























Z^£i 


&%> 


*7. 
























































































































icncH 
















1 — 






PUTflf 











































10 20 30 40 50 60 70 80 90 100 110 120 130 

Gallons per Mi'nu+e 

Fig. 285. — Performance curves of Maybach water pump. 

a reverse helical direction to the impeller. These passages 
connect with the common vertical outlet passage in the top of the 
body casting. The center portion of the top body casting, inside 



THE COOLING SYSTEM 375 

the helical passages above the impeller, is domed and fitted with 
a screwed plug. This plug is drilled with a small hole, to prevent 
an air-lock. Two other holes are also drilled in the bottom of 
the impeller between the vanes for the same purpose. The 
steel ball thrust race is exposed to the flow of water, a disadvan- 
tageous feature. Performance and efficiency curves for this 
pump at 2800 r.p.m. are given in Fig. 285. It will be seen that 
the maximum pump efficiency of 28 per cent is obtained with a 
discharge head of about 22 ft. of water and a capacity of 100 gal. 
per minute. 

Piping. — Water velocities in pipes vary from about 8 ft. per 
second in small engines to 16 ft. per second in large engines. 
Actual pipes sizes are from !}£ in. diameter for 90 h.p. to 2 in. 
diameter for 400 h.p. 

The frictional resistance to flow of water through straight 
pipes is given by /i = 4/ (l/d) (V 2 /2g) where h is the loss of head 
in feet, I and d are the length and diameter of the pipe respec- 
tively in feet, V the velocity in feet per second, and / is a coeffi- 
cient whose value is likely to vary from 0.004 to 0.010, depending 
on the roughness of the pipe. Inlets, outlets and bends will each 
offer a resistance equivalent to a length of 10 to 20 diameters. 

Large pipe sizes diminish the resistance and work of the 
pump but they weigh more and hold more water. The pump 
suction should be of ample diameter and as short and direct as 
possible. The connecting rubber hose should be firm and non- 
collapsible. Pipe lines should be of light tubing, bent to easy 
radii, with a minimum of bends and fittings. Hose connections 
at junction points should be very short, and should fit over cor- 
rugations. The fastenings should be by smoothly-bearing steel 
clamps which do not cut the rubber. Tape should be applied 
over hose and clamp and the whole shellacked. The pipes should 
be arranged to avoid air pockets if possible; if such occur, vent 
cocks must be applied. Particular care must be given to the vent 
cock on the pump casing. 

Water. — The water used should be free from lime. Filling the 
system with boiling water makes starting easy in cold weather. 
Anti-freezing solutions are all more or less objectionable, and 
it is best to drain the system when the plane is not in use. Fig. 
286 shows the properties of some anti-freezing mixtures. Alcohol 
lowers the boiling point and makes close control of temperature 
essential; the strength decreases and the freezing point is elevated 



376 



THE AIRPLANE ENGINE 



50 

45 

40 

£ 35 

3 
4- 
X 

E 30 



25 



20 









































<£ 

^ 


















x^ 


<- 


















V^ 








v<S- 
















V 












'V 

















































-10 -5 5 10 15 20 25 

Freezing Poiirf, Decj.Fahr. 

Fig. 286. — Properties of anti-freezing mixtures, 



30 



Water Tight 
Filter Cap 



Intake Manifold 




" Center Section 
To Expansion Tank 
Honeycomb / fromEngn 
Magneto 




•*. . Shutter Controf 
***•• Drain Cock 



• ' _ To Water Pump 

Water Pump _ , ... „ r 

„ , ' Front View of 



3^ ZVwz/7 Cfcetf 



Radi'a+or 



View 



Fig. 287. — Cooling plant of S E-5 airplane. 



THE COOLING SYSTEM 



377 



as the alcohol evaporates. Periodical tests by hydrometer are 
advisable. Glycerine does not evaporate, but impairs circula- 
tion and is detrimental to rubber. The glycerine should be 
stirred slowly into the water. 

Typical complete cooling systems are shown in Figs. 287 and 
288. Figure 287 is for a 180 h.p. Hispano-Suiza engine in a 
SE-5 plane with two tubular nose radiators at the sides of the 
engine, shaft, each 30 by 7^ by 3 1 %6 m -> with a total frontal area 
of 450 sq. in. and a radiating surface of 12,700 sq. in. The water 
capacity is 83J^ lb. and the flow rate 30 gal. per minute. The 
system is provided with an expansion tank which occupies the 



FillerCap 



To Motor 



Filler Cap . 
Vettf. 





— 4 

2 Separate Pipes X ^ ^ _ ___ y y 
from Radiator, to Pump ^- <"' 



Wafer Pump 



'•To Motor 
Top View of Rad/afor 



Fig. 288. — Cooling plant of Le Pere airplane. 



leading section of the middle panel of the upper wing. A small 
portion of the water leaving the cylinders passes around the 
intake manifold and is then returned to the pump; the rest of it 
goes through the radiator. The radiator is masked by shutters. 
Figure 288 shows a 360-h.p. Liberty engine in a Le Pere two- 
seater plane with a wing radiator in the center section of the 
middle panel of the upper wing. The radiator is 31 in. long, 27 
in. wide and 7 in. deep; has a frontal area of 783 sq. in. and a 
radiating surface of 35,520 sq. in. The water capacity is 41.6 lb. 
and the flow 80 gal. per minute Free water area 61.6 sq. in., free 
air area 1,247 sq. in., weight 127 lb. The water pumped around 
the manifold goes to the radiator before returning to the pump. 



CHAPTER XV 



GEARED PROPELLER DRIVES 



A well designed airplane engine develops its maximum power 
at a speed (r.p.m.) considerably in excess of the most efficient 
propeller speed. In order to combine maximum power develop- 
ment with most efficient utilization of that power it is necessary 
to resort to a geared drive. 

Geared drives have been employed in a number of successful 

installations. A German analy- 
sis of these 1 is the basis for the 
discussion which follows. The 
simplest type is a single-reduc- 
tion with spur gears as in Figs. 
289 and 290. In the Renault 
engine (Fig. 289), the gear ratio 
is two to one and consequently 
can be used for driving both 
camshaft and propeller shaft; 
in the Hispano-Suiza engine 
(Fig. 290) the gear ratio is four 
to three. Gears of this type 
show heavy wear. A design for 
a single reduction with internal 
gear is shown in Fig. 291; the 
internal gear housing is attached 
to the crankcase by an eccentric 
centering flange which permits 
accurate adjustment of the gears. This type permits great 
simplicity but there is difficulty in arranging satisfactory bear- 
ings on both sides of the gear wheels. 

With single-reduction gears the propeller shaft cannot be in 
the same axial line with the crankshaft; when this arrange- 
ment is desired double-reduction gears must be used. There 
are many possible arrangements; both pairs of wheels may be 
fitted with internal or external gears and in addition any one of 
the three shafts may be fixed while the other two drive and are 
driven respectively. Some of these arrangements are shown 
^utzbach: Technische Berichte, Vol. Ill, Sec. 3. 

378 




20T.P4.57t(inmm) 



Fig 



289. — Renault single-reduction- 
gear. 



GEARED PROPELLER DRIVES 



379 




Fig. 290. — Hispano-Suiza single-reduction-gear. 




Fig. 291. — Single-reduction internal gear. 



Crank Shaft and Propeller Turning in the Same Crank Shaft and Propeller Turning in 

Direction | Opposite Directions 

IntermediateShattTurnmg^ Intermediate Shaft 1 Turning m The Same 
in Opposite Direct/on 



Direction as CrankShaft 



Turning m The Same fcfa^MSfotf Turning 



in Opposite Direction 



im 



im 



h 



ft 



^ 



A 
Intermediate 

Shafting 
Fixed in 
Housing 



H 



jf 




JTt 



T-T 



Intermediate 
Shaft 
Revolvinq 
with * 
Propeller 




M 



& 



n t 



Intermediate 
Shaft 
Revolving 
with Crank 
Shaft 



Fig. 292. — Possible arrangements of double-reduction gears. 



380 



THE AIRPLANE ENGINE 



schematically in Fig. 292. In the top row the intermediate 
shaft is fixed; in the second and bottom rows it revolves forming 
the so-called planetary gears. The shaded gears are fixed and do 
not revolve. Some of these arrangements off er considerable diffi- 
culties for actual construction, notably in the matter of pro- 
viding suitable bearings on both sides of the gears; in others 
the space occupied may be great and the revolutions of the 
intermediate shaft very high. 

A simpler arrangement is one in which both pairs of gears 
have one gear in common. Schematic outlines of such reductions 
are shown in Figs. 293, 294 and 295. Figure 293 is developed 
from A t . Fig. 292; Fig. 294 from A 4 ; and Fig. 295 from B 2 . 
The Rolls-Royce planetary gear, Fig. 296, is an actual construc- 
tion of Fig. 295. The three revolving intermediate shafts are 




Fig. 293. Fig. 294. Fig. 295. 

Double-reduction gears with a common gear. 

carried in a spider, C. The internal gear, a, on the crankshaft 
drives the three gears, b, on the intermediate shafts, and the 
three gears, c, on the same shaft mesh with the gear, d, which is 
held against revolving in the housing. The spider, C, revolves 
and carries the propeller shaft. 

The advantage of the double-reduction gear over the much 
simpler single-reduction gear lies in the perfectly axial trans- 
mission of the power, from which the best condition of loading of 
the housing (pure torsion) is obtained. When the power is 
transmitted through two, three or four intermediate gears at equal 
angles, springing of the gear shafts from unequal peripheral forces 
or inaccurate tooth forms is avoided. Certain arrangements also 
make it possible to use heavy revolving masses (for instance, 
those of the intermediate shafts or the larger internal gears), 
thereby improving the uniformity of transmission and avoiding 
reversals of tooth pressure. The principal advantage, however, 
consists in the fact that on account of the load being divided 



GEARED PROPELLER DRIVES 



381 




382 



THE AIRPLANE ENGINE 



between two to four intermediate gears the tooth pressures per 
unit of tooth face are low. Consequently small pitches and small 
gears can be used which in turn have smaller construction 
defects since the defects of manufacture resulting from the use 
of inaccurate dividing wheels increase with increasing radius. 
The disadvantage of the double reduction gear is its weight and 
cost and the need for exact adjustment of the intermediate 
shafts if all the gears are to work equally. Furthermore, a 
complicated construction is necessary to ensure a solid and 
secure assembly of the gear. 

To obtain and keep proper adjustment of the reduction gearing 
as wear occurs, it is necessary to fit a joint between the crank case 
and the gear case, or, in the transmission, between crankshaft and 




Fig. 297. — Rolls-Royce single-reduction gear. 

gear, which will adjust itself automatically while running or 
can be adjusted in assembly. In the Rolls-Royce planetary 
gear, Fig. 296, a sliding cross linkage is used in a fixed housing. 
The link, (B) and e, lies between the outer engine housing and the 
intermediate gear wheel, d, which is held in the housing. Con- 
sequently the gear wheel, d, can adjust itself and always remains 
concentric with the crankshaft. The whole set of planetary 
gears also remains concentric with the crankshaft — which may 
shift in the casing — but not with the casing. The forward 
bearing, g and h, must be adjusted on each overhaul of the engine 
by the screws, /. 

In the Rolls-Royce spur gear, Fig. 297, the upper gear can 
be adjusted by eccentrically-set ball-bearing cages, c and d, and 
the lower gear can be adjusted on the engine shaft by screws. A 
universal joint is used between the crankshaft and the gear, a. 



GEARED PROPELLER DRIVES 383 

Many difficulties have been encountered in the actual oper- 
ation of reduction gears — principally, fracture, wear and heating 
of the gears. 

The bending stress in a gear tooth of the common involute 
form may be taken as 

W 
f = 14 X j- approximately, 

where W is the load in pounds on the gear tooth, b is its width 
and p is the circular pitch in inches. The mean value of the 
loading on the tooth can be determined from the known engine 
power, P, and the speed, V, of the pitch circle 

55^XP 
V 

The maximum loading on the teeth may be considerably 
greater than the mean loading either because of acceleration 
pressures, resulting from incorrect pitch or form of teeth, or 
because of irregular delivery of power from the engine, or on 
account of reinforced vibration near a resonance period of the 
shaft. The values of / calculated for a number of successful 
engines run from 30,000 to about 40,000 lb. per square inch; 
the material used is generally case-hardened chrome-nickel steel. 
These high stresses are calculated on the assumption that all 
the load is carried on one tooth. With accurate pitching the 
deformations of the loaded tooth will transfer load to the next 
tooth. With oblique teeth, such as herring-bone gears, the 
tooth pressure is distributed on an oblique line running from the 
root to the tip and the bending stress is thereby reduced. The 
stresses are worse if the teeth bear unevenly as a result of warping 
in hardening, untrue keying or poor forming. 

The surface pressure of the opposing curved tooth faces must 
not be sufficient to squeeze out the oil film. The relative sliding 
speed of straight-toothed gears is zero at the pitch circle, and 
the oil is more easily squeezed out under this condition than when 
there is relative motion. The bearing pressure is given by the 

W 

expression -7-7, where d is the diameter of the relative curvature of 

the teeth at the rolling circle. With involute teeth having radii 
of curvature of ei and e 2 at the rolling circle, 

2 _ 1 , 1 

-j — — ± —> 
a ei e<t 



384 THE AIRPLANE ENGINE 

the + sign applying to external gears, the — sign to internal 
gears. Calculations from successful engines indicate that with 
hardened gears the bearing pressure may go up to 1,400 lb. per 
square inch; if the gears are not hardened it should not exceed 
450 lb. per square inch. With internal gears the bearing pres- 
sures become low and hardening is, as a rule, unnecessary. 
Experience with roller bearings, where hardened rolls run between 
hardened rings, indicates a permissible bearing pressure of 2,800 
lb. per square inch or more at low peripheral speeds; if the rolls 
bear directly on the unhardened shaft the value falls to from 
150 to 300 lb. per square inch. 

With oblique toothed gears the contact shifts with great 
speed from side to side, as a result of which there is less tendency 
to squeeze out the lubricating film. 

Heating of the gears results from the sliding contact at the 
teeth and may be of such magnitude as to lead to trouble. The 
heat is best carried off by thermal conduction from the gears 
to the outer casing, but if this is not sufficient it must be assisted 
by oil cooling. The lubrication should not be so heavy that the 
oil heats up through churning; this may occur through the use of 
wide gears which catch the oil and force it out sideways with 
great force, or through locating the gears very close to the 
housing. 

For smooth running it is necessary that there should be no 
reversals of pressure in the gears. Four-cylinder engines give 
such reversals of pressure and so do six-cylinder engines at a low 
torque, or, with very heavy reciprocating parts, at high speeds. 
With a larger number of cylinders with crank angles equally 
spaced reversals will not occur. 

Central power plants have been used on several planes. The 
principal advantage which they offer is the possibility of con- 
centrating power plants in a central engine room (where they can 
be under constant supervision) and the resulting reduction of 
drag of the complete machine. There is also the possibility 
of reducing the number of mechanics required in a multi-engined 
plane. The disadvantages are the loss of power (possibly 5 per 
cent) in the transmission shaft and gears, and the increase in 
weight. 

Chain-driven propellers were used successfully by the Wright 
Brothers in 1903 and by others later. In recent years the chain 
drive has not been used but shaft and bevel gears have been 



GEARED PROPELLER DRIVES 385 

employed with some degree of success. Siemens-Shuckert 
multi-engine planes of several sizes have used bevel gear drives. 
The largest of these with six engines and four propellers, is ar- 
ranged with the four rear engines driving the two rear propellers 
at half engine speed and the two front engines driving the two 
front propellers with a reduction ratio of 14 to 9. The couplings 
between the engines on the main transmission gear are a combina- 
tion of friction and independent couplings. The latter enable the 
engine to be disengaged and stopped if damaged. The articu- 
lated transmission shafts are connected at both ends through 
laminated spring couplings. 

The main difficulty in the operation of shaft drives has been in 
the setting up of " torsional resonance," which has caused break- 
age of shafts and universal joints. This has been overcome 
by the use of a flywheel on the engine and a special clutch which 
combines a dog clutch and a friction clutch. The shaft should 
rotate at engine speed and the gear reduction should be near the 
propeller. Some trouble has resulted from " whirling" of long 
shafts but probably because bearings have been placed at nodal 
points; this can be avoided. With these difficulties overcome 
it is doubtful whether the expense, weight and complication of 
the flywheels, clutches, shafts and gears will not more than 
counterbalance the advantages of a central engine room. 



25 



CHAPTER XVI 
SUPERCHARGING 

Change of Engine Power with Altitude. — The indicated work 
in the cylinder of a gasoline engine is the product of the heat of 
combustion of the fuel by the thermodynamic efficiency of the 
engine. The thermodynamic efficiency is unaffected by the air 
density and depends only on the ratio of compression. The heat 
of combustion is determined by the weight of fuel which can be 
burned and this depends on the weight of air admitted and con- 
sequently on the density of the air. All other conditions remain- 
ing constant, the indicated power of an engine would vary directly 
as the density of the air. 

The brake horse power of the engine is the difference between 
its indicated power and the power required to overcome engine 
friction. At constant engine speed the friction will not change 
greatly with the air density; it increases with lowered temperature 
of the lubricant and decreases with lowered pressures at rubbing 
surfaces. If the frictional resistance remained constant the 
brake horse power would fall off much more rapidly than the 
indicated power at high altitudes. For example, an engine 
developing 100 i.h.p. at the ground will give 85 b.h.p. with 
15 friction h.p. Operating in air at one-half ground density 
the theoretical indicated power is 50 h.p. and with 15-h.p. friction 
there would be 35 b.h.p. The brake power would be diminished 
in the ratio 35/85 = 0.412, while the indicated power is halved. 

The actual diminution in brake power is not as great as the 
preceding calculation would indicate; the conditions are complex 
and not susceptible of exact calculation. 

The friction horse power is partly rubbing friction and water 
and oil pump work and partly work done in overcoming throttling 
losses at the intake and exhaust of the gases. The former losses 
may be assumed constant with varying air density; the latter may 
be assumed to vary directly as the air density. If the total 
fricton loss is 15 per cent of the full indicated power at the 
ground, and the throttling losses are assumed to be one-third of 
the total friction loss, and if the indicated horse power is propor- 

386 



SUPERCHARGING 



387 



tional to the relative air density, d, then the brake horse power, B, 
at any air density is given by 

D d - 0.10 - 0.05d D 0.95d - 0.10 D 

t> = T^TZ Bo = TTTT^ -Do 



0.85 



0.85 



where B is the brake horse power at the ground. The following 
table gives horse powers calculated for different altitudes. It 
will be seen that the brake horse power is nearly proportional to 





Relative 


Relative 


Relative 


Altitude, feet 


air density, d 


air pressure 


b.h.p., B/B 





1.0 


1.0 


1.0 


6,000 


0.829 


0.801 


0.808 


12,000 


0.694 


0.645 


0.660 


18,000 


0.581 


0.518 


0.532 


24,000 


0.485 


0.414 


0.424 


30,000 


0.411 


0.333 


0.342 



the air pressure. Actual tests support these calculated quantities 
for altitudes up to 10,000 ft.; for high altitudes the brake horse 
power does not decrease as rapidly as the air pressure nor so 
slowly as the air density but according to some intermediate law. 

Tests made at the Bureau of Standards 1 show a variation of 
brake horse power with barometric pressure as in column 7 of 
Table 20. The ratio of brake power to air density is given in the 
eighth column. It is seen that the brake power falls off more 
rapidly than the air density and that at one-half ground density 
the brake power is about 0.43 time the ground power. 

The variations with air density of the mechanical, volumetric 
and thermal efficiencies of the Liberty 12 engine and the Hispano- 
Suiza 300 at a speed of 1,600 r.p.m. are given in Fig. 298. 

The relative horse powers of Table 20 are based on constant 
engine speed. They may more properly be regarded as relative 
engine torques. Engine speed falls off with increasing altitude 
so that the actual horse power developed falls off more rapidly 
than is indicated in Table 20. With constant revolutions per 
minute the resisting torque at the propeller diminishes in direct 
proportion to the air density and consequently falls off less 

1 4th Annual Report, National Advisory Committee for Aeronautics, 1918, 
p. 502, Fig. 6. 



388 



THE AIRPLANE ENGINE 



rapidly than the engine torque. Since the engine torque is 
practically independent of the revolutions per minute the engine 
speed will diminish as altitude is gained until that speed is reached 
at which propeller torque equals engine torque. 

The actual engine power at any altitude is given by 

N 



P =P G XKX 



N G 



where P G is power developed at the ground, 
Nq is revolutions per minute at the ground, 
N is revolutions per minute at altitude, 
K is the quantity in the seventh column of Table 20. 



90 



+ 90 
J 30 
J> 70 




Barome+n'c Pressure in Cm. of Hg. Approximate 



J M 



^L 



90 
80 
70 

+ 100 
v§90 
£80 

30 
Z0 
10 



74.51 


63.29 


52.51 - 


43.43 


35.81 29.74 






! i 




i 








i 




Mechan, 


fZES&k 
















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1800 Rpm. 


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0.030 0.070 0.060 0.050 0.040 0.030 O.OoO 0.070 0.060 0.050 0.040 0.030 
Air Densitx) in Lb. per Cu. Fh 
Liberty 12. Hispano-Suiza 300. 

Fig. 298. — Variation of engine efficiencies with air density. 



Table 20 shows that the engine power at constant speed is almost 
exactly proportional to the barometric pressure. On this basis 
the engine power at an elevation where the barometer is B cm. is 

f - f G x 76 x Ng 

Supercharging. — The diminution in power of a gasoline engine 
with increasing altitude results in a moderate reduction of speed 
in horizontal flight. If greater power were available the ground 
speed could be maintained at all elevations or exceeded, if desired. 
Much effort has been expended in attempts to prevent or reduce 



SUPERCHARGING 



389 



03 


in 
in 

Q. 


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Brake horse 
power devel- 
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Air tempera- 
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390 THE AIRPLANE ENGINE 

this falliDg off in engine power. Such falling off would be entirely 
avoided if steam power could be substituted for gasoline power, 
since the boiler and condenser pressures would be independent 
of the barometer pressure. Attempts to design a light-weight 
steam plant have not been successful; there is no difficulty with 
engine or condenser (which takes the place of radiator), but it 
has not been found practicable to design a boiler to withstand 
high steam pressures and of sufficiently extended heating surface 
without arriving at weights which are prohibitive for airplane use. 
Furthermore, the lower fuel economy of a steam plant would 
necessitate the carrying of a greater weight of fuel. 

Two general methods present themselves for increasing gas 
engine power at high altitude : 

1. To select an engine so large that it will give the desired power when 
running with wide-open throttle at the high altitude at which the airplane is 
intended to fly, and to operate it at partial throttle at all lower altitudes. 

2. To select an engine which gives the desired power at the ground and 
add some device for supplying the cylinder with air at a pressure greater than 
the barometric pressure when desired. This process is known as pre- 
compression or supercharging. 

As an illustration, suppose it is desired to fly at 20,000 ft. 
developing 400 h.p. This can be accomplished either by install- 
ing an engine which would develop 800 h.p. with wide-open 
throttle at the ground; or by installing a 400-h.p. engine provided 
with a supercharging device which is able to maintain that horse 
power at all altitudes up to 20,000 ft. If the large engine is used 
the engine weight will be increased. An estimate made of the 
in crease of weight which would result from doubling the power of 
a Liberty motor by doubling the piston area per cylinder, while 
keeping the stroke constant, indicates this increase would be about 
40 per cent. If the power were doubled by doubling the number 
of cylinders the weight would be nearly doubled. It should be 
noted that if the engine is not permitted to develop more than 
400 h.p. at any elevation, the radiator, water pump and general 
cooling system will not be larger than for a 400-h.p. engine. If 
the smaller engine is used the engine weight will also be increased 
by the addition of the supercharging apparatus and the engine 
becomes more complicated. 

Oversized Engine. — In this system, the greater weight of 
the engine is offset not only by greater simplicity (as compared 
with a supercharging engine) but also by greater economy. Such 



SUPERCHARGING 



391 



engines should be provided with an automatically controlled 
throttle valve, actuated by some device (generally similar to an 
aneroid barometer) which responds to changes in atmospheric 
pressure. An example of such a device is given in Fig. 299 in 
which an airtight flexible chamber filled with air at low pressure 
actuates a balanced double-seated throttle valve. If the throttle 
is placed before the carburetor, the top of the float chamber 
must be kept in communication with the low-pressure side of the 
throttle. The control may be so adjusted as to give constant 
horse power at all altitudes up to that at which the throttle is 
wide open; the power cannot be maintained beyond that point. 
With an engine so operated it is possible to use a higher ratio 
of compression, without danger of preignition, than with an 
engine which has wide-open throttle at the ground. 




Fig. 299. 



. Ouiside Air In fer 
XJ and Inlet from Compressor 
-Automatic throttle control for oversized engine. 



With constant power output the weight of the charge admitted 
per cycle will be approximately constant and the pressure in the 
cylinder at the beginning of compression is also approximately 
constant. The latter quantity is actually a little more at the 
ground than at higher altitudes because the engine is exhausting 
against a higher barometric pressure and consequently there is a 
greater weight and pressure of burned gases remaining in the 
cylinder to be mixed with the new incoming charge. Further- 
more, the efficiency at the ground would be lower than at high 
levels on account of the higher back pressure. With constant 
power output the pressure in the cylinder at the beginning of com- 
pression would be less than 7 lb. per square inch at 20,000 ft. 
elevation, and probably less than 8 lb. per square inch at the 
ground. That is, the maximum pressure to be expected at the 
beginning of compression is 8 lb. per square inch as compared with 
14 lb. in the usual engine. This results in lower compression and 



392 



THE AIRPLANE ENGINE 



explosion pressures. Furthermore, the cylinder temperatures 
are lower throughout the cycle mainly in consequence of the 
smaller amount of heat developed by the explosion and the 
greater cooling effect of the water jacket. Under these conditions 
it is possible to employ higher compresssion without danger of 



I 



Meters 
4627 



1 8. 

° jS I0S[- 
o 3: c 

+ £ Y 

8-M 

'2 j I 

1 ih E 
o £ <3 
■- g lb 
tc uj b 90 





















. 


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C-CompJ_ ^""i "-[ T" 
























S-Compression Ratio 5-3 














































































° 












r i ii " - 










































































c 


?c 


00 


6 


00 


10 


000 


14 


300 


18,000 


22JJ0O 


26000 


30,000 



Altitude, Feet. 

Fig. 300. — EffectTof altitude on variation of engine power with compression 

ratio. 



preignition. Engines have been operated in this manner with a 
ratio of compression as high as 7. 

The employment of a high ratio of compression will increase 
the available power, particularly at high altitude. This is shown 
clearly in Fig. 300, which gives the results obtained at the Bureau 
of Standards with an engine supplied with three different sets of 
pistons to give different ratios of compression. The curve B is for 
a compression ratio of 5.3, which is here regarded as standard. 
The curves A and C are for compression ratios for 4.7 and 6.2 
respectively. It will be seen that, calling the horse power with 



!3>5 



T ,25 

D. 

JS 120 













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1 100 1200 1300 1400 1500 1600 
R.P.M. 



5.5 5.65 5.6 
Compression Ratio 



Fig. 301. — Variation of engine capacity with compression ratio and engine 

anooA 



standard compression 100, at all altitudes, it is increased to 104^^ 
at 20,000 ft. with the high compression and reduced to 95Ji 
with the low compression. At the ground the corresponding horse 
powers are 102J4 and 96%. German tests on a Benz 200-h.p. 
engine, Fig. 301, show similar increase of power with ratio of com- 



SUPERCHARGING 



393 



pression and show also that the actual increase may be greater 
than the theoretical (air c} T cle) increase, particularly at high 
engine speed. 

The increase in engine size necessary to maintain ground 
power is inversely as the density of the air at the altitude up to 
which full power is desired. The percentage increase is shown in 
Fig. 302. 

The change in horse power actually developed with varying 
compression will be greater than these constant-speed tests 
indicate. With a given propeller the engine speed will increase 



200 


1 


1 




> 


180 


1 










^x*^ 












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5 IW 
















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120 


























1 


100 






1 









4000 



:e::? 



20000 



Fig. 302. 



&000 12000 

Al+i+ude. Fee+ 
-Required displacement volume of oversized engine. 



with the engine torque and will cause a further increase in horse 
power. French tests show the following results: 



Hispano-Suiza 



LeRhone 



Clerget 



Ratio of compression ■ 5.3 

Revolutions per minute ; 2 , 040 

Brake horse power 186 



5.8 

2,070 

195 



5.18 
1,230 
119 



5.65 
1,260 
126 



6.58 
1,290 
134 ' 



4.6 
1,290 
123 



5.2 
1,350 

137 



5.6 
1,360 
144 



The admission of inert gases with the explosive mixture is 
now being developed as a means of maintaining high economy 
at low levels with an oversized engine. The inert gas is cooled 
exhaust gas. Its presence will permit operation with full 
throttle at much higher compression ratios than would otherwise 
be possible and consequently with higher thermal efficiency. 
As elevation is gained the percentage of inert gas in the charge 
may be reduced either by hand control or automatically; it should 
be possible to maintain the engine power at high altitudes and 
to operate continuously at very high efficiency by this device, 
(see p. 435). 



394 



THE AIRPLANE ENGINE 



Up to 12,000 ft. altitude the oversized engine (with about 
46 per cent increase in volume) is preferable both in respect of 
total weight and of simplicity of construction to the supercharged 
engine. With altitudes in excess of 20,000 ft. the weight of the 
oversized engine becomes considerable and the preference may 
rightly fall on the supercharged engine. A combination of the 
two may possibly turn out to be best, with power maintained 
constant up to about 10,000 ft. by the gradual opening of the 
throttle valve, and then bringing into action a supercharging 
device to maintain constant power up to 20,000 ft. 

Supercharging Engine. — In a supercharging engine, air is 
compressed by a blower or other device and is delivered to the 



1.15 

1.10 
1.05 

1.00 

095 

0.90 

'■"- 0.85 

2 0.80 

fc 0.7S ■ 

o 

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o 

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75 70 65 60 55 50 45 40 55 30 25 20 15 
Exhaust Back Pressure (Cms.ofHg.) 

Fig. 303. — Chart for finding the power developed by a supercharged engine. 



carburetor at a pressure in excess of the surrounding atmospheric 
pressure and consequently in excess of the exhaust pressure, 
except in the case (discussed later) where the blower is driven by 
an exhaust gas turbine. The power that can be delivered by 
an engine whose admission and exhaust pressures are different is 
readily calculable for an ideal engine, but it is necessary to have 
recourse to actual tests in order to ascertain its magnitude for an 
actual engine. Such tests have been conducted at the Bureau 
of Standards; the curves given in Fig. 303 show the results 
obtained. These curves give the horse power that will be 
developed by an engine with any exhaust pressure from 76 to 20 
cm. of mercury and with air supplied to the carburetor at any 
pressure from 76 cm. down to 55 cm. of mercury. The horse 



SUPERCHARGING 395 

power is given as a ratio to the horse power delivered at the ground 
with admission and exhaust both at a pressure of 76 cm. of mer- 
cury. The tests were conducted with the carburetor adjusted to 
give maximum power and the curves are all corrected to the same 
air temperature at the carburetor. The curves show that if the 
pressure at the carburetor is maintained at 76 cm. during flight 
the horse power developed in the engine will increase as a result 
of diminishing back pressure and at 20,000 ft. (36.5 cm. pressure) 
will be about 6 per cent greater than the horse power at the 
ground; this is to be compared with the diminution of 51 per 
cent in horse power (see Table 20) at the same altitude without 
supercharging. 

The gain in engine horse power with supercharging is not of 
course net gain. Some of the additional work is used up in pre- 
compressing the air to the admission pressure. Furthermore, the 
precompression heats up the air so that it enters the carburetor 
at a temperature greater than that of the surrounding atmosphere. 
The actual horse power developed in the cylinder is given by the 
equation 

Pc = Po X r X F 

where P c is the horse power developed with the supercharging 
apparatus at the given altitudes; P G is the observed horse power 
on the ground at the observed carburetor air temperature, fa; r is 
the horse power ratio at the given condition of exhaust and 
carburetor pressures produced by the supercharging device at the 
given altitude, (obtained from curves, Fig. 303) andF is the temper- 
ature correction factor to correct from observed temperature 
at the ground, fa, to temperature at the carburetor, t 2 . 

The temperature of the precompressed air can be calculated 
from the equation 

n-l 

T 2 






where T 2 is the absolute temperature of the air entering the 
carburetor, T s is the absolute temperature of the air entering the 
supercharging device, p 2 and p 3 are the air pressures at the same 
places. The quantity n may be assumed for ordinary conditions 

T 

to have the value 1.3. The quantity ~r may be obtained from 

J- 3 

Fig. 304, which gives values also for n = 1.2 and n = 1.4. 



396 



THE AIRPLANE ENGINE 



The effect of this increase of temperature is to diminish the 
power of the engine. The temperature correction factor is 
obtained from the equation (compare p. 33). 

_ 920 + h 
920 + t 2 
where t\ and t 2 are in degrees Fahrenheit (not absolute). 

As an example of the use of the above, suppose an engine 
develops 200 h.p. at the ground with barometer 76 cm. mercury 
and temperature 40°F. and that it is taken to an altitude where 
the barometer is 35 cm. and temperature — 4°F. and that it is 
supercharged to 65 cm. pressure. From Fig. 303 the horse power 

1.3 



£lf* 



1.0 I.I 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 21 22 2.3 24 25 
Fig. 304. — Temperature rise of air during compression. 

















* 


































^> 
























^ 




*>r 



































































































,. Vi 65 
ratio — = ^-^ 
Ps 36.5 

T 2 



ratio, r, corresponding to these conditions is 0.9. The pressure 
= 1.78. Assuming that n = 1.3, Fig. 304 gives 

1.142, or T 2 = 1.142 X (460 - 4) = 520. Consequently 

J- 3 

t 2 = T 2 — 460 = 60°. The temperature correction factor F = 

Q20 4- fiO = 0-978- The engine horse power will then be 

Pc = 200 X 0.9 X 0.978 = 176 

If the barometric pressure at which the ground horse power is 
observed is not 76 cm., a further correction may be introduced. 
For example, if the barometer reads 74 cm., the horse power ratio 
as compared with 76 cm. is found from curve E, Fig. 303, to be 
0.972. The horse power actually developed, P c , will then be 
176 



0.972 



= 181. 



As previously pointed out, this horse power is not the net 
horse power available for driving the propeller. There must be 
subtracted from it the work required to precompress the air. 



SUPERCHARGING 



397 



The ideal method of compression is isothermal but this cannot be 
realized. If there were no addition or abstraction of heat 
during the compression and no frictional resistance or eddy losses, 
the compression would be adiabatic and this is what the cen- 
trifugal compressor, without cooling, might be expected to accom- 
plish. The work of adiabatic compression is given by 



W 



= wR 



(p&i — P3V3) 
(T 2 - T z ) 



7- 1 
= wJ C p (T, - T 3 ) 

where p 2 , P3, are the pressures at beginning and end of compres- 
sion respectively; v 2 , v i} the corresponding volumes; T 2 , T 3 , the 



55 

g45 
w 40. o 
£.35° 
■j=30 2 

f> 20 £ 

fe 15°" 

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3.2 

3.0 
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p * 


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2.0 
1.8 










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1.2 
10 


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r 















too e 

90 I" 
80 "3 

70 |h 

60 ^^r 
50 1? 1 
40 *£ 
30 £ Q 
20 % 
10 £" 



1200O 16000 
Altitude m Feet 



20O00 24000 



Fig. 305. — Power absorbed in the compression of air. 

corresponding absolute temperatures; R is the air constant, 53.4; 
J the mechanical equivalent, 778; C p the specific heat of air at 
constant pressure, 0.241; 7 the ratio of specific heats, 1.4; and 
w the weight of air compressed. The final temperature, T 3 , 

is given by the equation 

7-1 

T 2 \pj 

The work done per pound of air compressed (adiabatically) per 
second is given in Fig. 305, curve W, for compression from the 
mean pressure and temperature existing at any altitude to 
standard atmospheric pressure. This quantity is expressed in 
curve P as a percentage of the work which can be done by that 
air in an efficient engine. The ratio of compression (pressure 
ratio) is shown by curve C and the temperature rise of the air by 
T. The mean temperatures used as a basis for calculating the 
above curves are given by Fig. 275. 



398 THE AIRPLANE ENGINE 

With isothermal compression the work required to compress 

1 lb. of air is given by W = RT 3 log e — ft.-lb. The actual work 

done on a centrifugal compressor in compressing air is found to be 
about twice the work of isothermal compression. For the exam- 
ple worked out above, the work of isothermal compression for 

1 lb. of air is W = 534 X 456 X log e 1.78 = 14,040 ft.-lb. 
The actual work of air compression per pound of air will be 

2 X 14,040 = 28,080 ft.-lb. 

The amount of work done by 1 lb. of air in the cylinder is 
determinable from the assumption that the explosive mixture is 
15 parts air to 1 part gasoline, by weight, and that the fuel 
consumption is % lb. gasoline per horse power hour. Every 

A « ' A 1 2 X 33 > 000 X 60 OKA AAA *x 1U r* 

pound of air does work ~ = 264,000 ft.-lb. Of 

28 080 
this work the fraction ^ ' ~„~ = 0.1065 is used for precompres- 

sion. Consequently in the case discussed the net available 

horse power will be 181 X (1 - 0.1065) = 161 h.p., that is, 

20 h.p. will be used in driving the blower or other supercharging 

device. 

There is one type of supercharging device in which the power 

required for precompressing the air is obtained from an exhaust 

gas turbine. This imposes a back pressure during the exhaust 

stroke in excess of the atmospheric pressure. For example, 

suppose that under the same conditions as those worked out in 

the example with 65 cm. carburetor pressure, an exhaust gas 

turbine is used and that the exhaust back pressure is 60 cm. 

mercury. The horse power ratio is now (Fig. 303) 0.85 and the 

85 
net horse power is -~^ X 181 = 171 h.p. 

In the exhaust gas turbo -superchargers that have been built 
up to the present, the selected operating conditions have generally 
been the maintenance of ground pressure in both admission and 
exhaust manifolds up to some limiting altitude. Assume a 
ground pressure of 76 cm., temperature 66°F. and the engine 
operating at an altitude where the barometer is 38 cm. (19,000 
ft.), and air temperature 5°F. If the exponent n during the - 
compression has. the value 1.3, it is seen from Fig. 304 that 

=r = L174 for ^ = 2. As T 3 - 460 + 5, T 2 = 545, or the 

I 3 ??3 



i 



SUPERCHARGING 399 

temperature of the air entering the carburetor is 545 — 460 = 
85°F. The engine horse power is thereby diminished in the 

The work W available from the exhaust gas turbine may be 

determined from the equation for adiabatic compression given 

above. The velocity, V, with which the gas discharges on the 

blades of the turbine (assuming a frictionless nozzle) is given by 

V 2 
the equation W = «-, or it may be obtained from the equation 

V = yJ2g -^-j RT 2 \ 1 - (^- 3 ) 7 I ft. per second. 

where T 2 is the absolute temperature and p 2 the absolute pressure of 
the exhaust gases entering the turbine nozzle. Values of V from this 
equation are given in following table calculated for T 2 = 1,800. 
The average temperature of the gas leaving the engine is about 
1,500°F., but loss from the exhaust manifold reduces it to about 
1,300°R, or 1,800° absolute. 



1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9, 2.0 
Vz 

V 751 1.026 1,230 1,384 1,514 1,622 1,716 1,798 1,871 1,940 
' 

With an exhaust gas turbine of 100 per cent efficiency, the work 
that can be done by the gas is equal to the kinetic energy of the 
gas. If an exhaust gas turbine, maintaining ground pressure at 
the carburetor and in the exhaust manifold, is fitted to a 200-h.p. 
engine, using 0.5 lb. gasoline per horse power hour and 15 lb. 
of air per pound of gasoline, the weight of exhaust gases will be 

200 X H Q ° X 16 = 26.7 lb. per minute. At 19,000 ft, altitude 

p 2 
(38 cm. barometer) the pressure drop ratio — in the expansion 

nozzle is 2, and the corresponding gas velocity is 1,940 ft. per 

26.7 X 1 940 2 
second. The kinetic energy of the gas is — ^ — ' — = 26,000 

ft.-lb. per second or J- n = 47.3 h.p. Assuming an efficiency 

of 50 per cent for the gas turbine the power available for driving 
the compressor is 23.7 h.p. The compressor work with isother- 
mal compression would be W = 534 X 465 X log* 2 = 17,400 



400 THE AIRPLANE ENGINE 

ft.-lb. per pound of air compressed. The weight of air com- 

, . 200 X 0.5 X 15 A -._v, , _ 

pressed is ArTw^o = 0.417 lb. per second. The power 

required for isothermal compression is — ,, ' =13.2 

h.p. If the compressor efficiency is 50 per cent as compared 

with isothermal compression, the power required to drive the 

compressor will be 2 X 13.2 = 26.4 h.p. 

For the operation of the turbo compressor just discussed, 

it is necessary that the over -all efficiency of the combination 

13.2 
should be not less than j^q = 0.279. This over-all efficiency, E, 

is the product of the turbine efficiency E T and the compressor 
(isothermal) efficiency E c , or E = E T X E c . Tests on the 
Rateau exhaust gas turbo-compressor indicate the possibility of 
values of E T of about 0.53, and a value of about 0.5 for E c . 
These correspond to E = 0.53 X 0.5 = 0.265. The theoretical 
work of isothermal compression which can be done by this 
combination for the special case under discussion is 47.3 X 0.265 
= 12.55 h.p. or is less than the 13.2 h.p. calculated as necessary 
to maintain ground pressure at the carburetor. 

The actual pressure which could be maintained at the carbu- 
retor is readily calculable. The work of isothermal compression 

per pound of air is W = RT 3 log e — > and as the weight of air is 

0.417 lb. per second, the horse power for isothermal compression 

is (0.417 X RT 3 log e -)/550 = 12.55. The value of T 3 has been 

given as 456. Solving the equation gives — = 1.95, and pi = 

Pz 

74 cm. That is, the power developed by the exhaust gas turbine 

is sufficient to compress the air to 74 cm. pressure. If a higher 

pressure is desired at the carburetor either the back pressure 

on the engine must be increased, thereby increasing the turbine 

power, or the efficiencies of turbine and compressor must be 

increased. 

The inefficiency of the compressor has a further effect on 

the engine performance besides that just discussed. Practically 

all the work done on the compressor goes finally into heating of 

the air and as the (isothermal) efficiency of the compressor is 

about 0.5 the amount of such heating can be readily determined. 



SUPERCHARGING 401 

For the case under discussion with air at 38 cm. pressure, and 
465° absolute temperature Fahrenheit and with compression to 
74 cm. pressure, the work of isothermal compression per pound 
of air is 53.4 X 465 X log e 1.95 = 16,600 ft.-lb. = 21.35 B.t.u. 
The total work done is 2 X 21.35 = 42.7 B.t.u. and the conse- 

42.7 4.2 

quent heating of the air is = OA * = 177°F. The final 

temperature of the air will be 177 + 465 = 642° absolute = 182°F. 

This high temperature of the air entering the carburetor will 
cause a decrease in volumetric efficiency of the engine. To avoid 
the consequent loss of engine power the air should be partly cooled 
on its way from the compressor to the engine. Some heating of 
the air is highly advantageous in aiding the vaporization of the 
fuel in the carburetor and intake manifold. 

Centrifugal compressors (single stage) do not appear to be 
suitable for ratios of compression greater than 2 to 1 on account 
of the excessive speeds which become necessary. This means 
that they can be used only up to altitudes of about 20,000 ft. if 
they are to maintain ground pressure at the carburetor. Multi- 
stage compressors permit higher ratios of compression, or the 
same ratio with lower peripheral speeds. 

Supercharging Devices. — Two methods have been employed 
for supplying the engine with air at a pressure higher than that of 
the surrounding atmosphere. 

1. The cylinder takes in an overrich charge in the usual 
manner and this charge is raised in pressure and diluted to the 
proper strength by the admission of compressed air at the end of 
the admission stroke. 

2. The whole of the air going to the engine is compressed 
and is sent under pressure through the carburetor. 

In the first method the demand for compressed air is intermit- 
tent and the compression is most suitably carried out in a recip- 
rocating (piston) compressor. The second method requires a 
centrifugal compressor. It suffers the disadvantage that the 
pressure in the carburetor is greater than the external pressure 
so that the carburetor must either be made strong and tight 
enough to withstand this condition or must be entirely enclosed 
in a chamber under the compressor pressure, which makes it 
comparatively inaccessible. 

An example of the first method is the Ricardo system, which 
has been experimented with considerably in England. The 

26 



402 



THE AIRPLANE ENGINE 



Hand Operated 
Valve Here 
IF 



Infer Cooler 



cylinder (Fig. 306) has a ring of ports uncovered near the lower 
end of the stroke. The piston is of two diameters, as shown, 
leaving an annular space, A, which diminishes in volume as the 

piston descends. This annular 
space communicates freely with 
the intercooler, B, which con- 
nects with the ring of ports. 
The closing of a hand-operated 
valve at F puts the super- 
charger out of action when 
desired. Air is admitted to 
the annular chamber, A, 
through the automatic valve, 
E. 

In Fig. 306 the piston is 
shown near the end of the 
suction stroke. Compressed 
air from B is just beginning to 
flow into the cylinder through 
the ring of ports; the inlet 
valve is nearly closed and the 
pressure in the cylinder is 
raised. During the succeeding 
compression stroke air is ad- 
mitted through E into A and 
is compressed in A and B dur- 
ing the expansion stroke . Near 
the end of the expansion stroke 
the ring of ports opens and some 
of the burned gases pass into B ; 
the exhaust valve opens im- 
mediately afterwards and these 
the compressed air sweep back 
A new charge of 




Automatic Air Valve 



Fig. 



306. — Ricardo supercharging 
gine. 



burned gases together with 
through the ports and scavenge the cylinder 
air is taken in through E during the exhaust stroke and is com- 
pressed during the following suction stroke. 

It is obvious that this system can be used only for moderate 
degrees of supercharging since the additional air supplied to the 
cylinder per admission cannot be greater than 



L X AD* - d*) 



SUPERCHARGING 



403 



where L is the piston stroke, and D and d are the two pis- 
ton diameters. As the normal cylinder charge has a volume 

D 2 -d 2 /Hx 2 



L X jD 2 this represents an increase of 



Z> 



= 1 



o 



If d = J^D, the increase in charge would be 75 per cent. The 
results of tests with this device are given in Fig. 307; they show 
an increase of power of approximately 50 per cent. 

The more promising method of supercharging appears to be 
that in which all the air going to the engine is precompressed in 
a compressor. Reciprocating compressors operating at engine 
speed have been tried in England but show an over-all efficiency 
which is very low — only about 21 per cent. The Roots type of 
positive blower (Fig. 308) with aluminum rotors operating at 



100 






.6. j/ 






b 80 

<: 
o 

0. 

5? 60 
S 40 

CD 

?0 




cMS A 


^ 






1$> 


3^ 

















800 



1000 



1400 



1200 
R.pjn. 

Fig. 307. — Performance curves of 
Ricardo supercharging engine. 




Fig. 308.— Roots blower. 



twice engine speed gives an over-all efficiency of about 52 per 
cent but is exceedingly noisy. The most commonly used type 
is the centrifugal compressor. Such a compressor may either be 
directly coupled to the engine or driven by a separate engine 
or by an exhaust gas turbine. In any case, its peripheral speed 
must be high in order to keep down the number of stages and the 
weight and bulk of the compressor. If directly coupled to 
the engine shaft a train of gears must be employed to increase the 
speed up to 10,000 r.p.m. or more, depending on the number of 
stages employed and the amount of supercharging desired; if 
operated through a gas turbine no gears are necessary as the 
turbine speed will be from 20,000 to 30,000 r.p.m. and one 
compressor stage will be sufficient at these speeds. 

Geared direct-coupled compressors have given much trouble 
from stripping of gear teeth. At the high speeds of rotation 
required, the kinetic energy of the rotor wheels is very great and 



404 



THE AIRPLANE ENGINE 



considerable forces have to be employed for rapid acceleration. 
When the engine is started or the throttle valve is opened sudden- 
ly the pressure at the gear teeth is so high and so suddenly 
applied that breakage is likely to occur. To prevent this a 
friction clutch, spring coupling, centrifugal clutch or other 
equivalent device must be employed between the engine shaft 
and the rotor wheels. 

One solution of this problem is shown in the Sturtevant super- 
charger, Fig. 309. The single-stage blower runs at 10 times 
the engine speed through a 2 to 1 belt drive in series with a 5 to 
1 helical gear drive contained within the blower casing. The belt 
drive is vertical and is brought into action when desired by an 

Induction Pipes 




Fig. 309. — Sturtevant supercharger. 

idler pulley. This arrangement gives ample slip when the 
engine speed charges suddenly. The weight of the super- 
charging device in this case is stated by the manufacturers to be 
50 lb. for a 210-h.p. engine. The engine speed with constant- 
pitch propeller increases from about 2,100 r.p.m. at the ground 
to about 2,500 r.p.m. at 20,000 ft. altitude; the blower 
consequently increases from 21,000 to 25,000 r.p.m. 

An English design, shown in Fig. 310, has a double reduction 
of 11 to 1 between the engine shaft and the blower disc with 
three intermediate wheels distributing the torque to the driven 
pinion; the over-all efficiency is 53 per cent. 

German constructions show multi-stage compressors with the 
relatively low peripheral speeds of 400 to 500 ft. per second. 1 
At these speeds the design works out to three stages to maintain 

1 Hildesheim, Automotive Industries, Oct. 21, 1920. 



SUPERCHARGING 



405 




406 



THE AIRPLANE ENGINE 



full power up to 11,500 ft. and four stages to 16,000 ft. The 
compressor makes 10,000 to 11,000 r.p.m. Such compressors 
have been used both as individual superchargers direct-coupled 
to a single engine, and as central superchargers driven by a 
separate engine and delivering compressed air to all the engines 
of a multi-engine plane. With a separate engine, no clutch is 
necessary between engine and compressor but the engine must 
be provided with a flywheel to prevent too great acceleration 
and consequent stripping of the gear wheels. Individual super- 
chargers are generally driven from the rear end of the crankshaft, 




Fig. 311. — Schwade multi-stage centrifugal supercharger. 



but in some cases the gears connect to the propeller end of the 
shaft to avoid the torsional oscillations which have sometimes 
given much trouble at the rear end. 

The Schwade three-stage supercharger shown in Fig. 311 
delivers 2,200 lb. of air per hour at a pressure ratio of 1 to 1.52 
(11,500 ft. altitude). The shaft speed is 1,400, intermediate 
gears 3,500, blower 10,500 r.p.m. The rotor diameter is 10 in., 
peripheral speed 460 ft. per second. The pinion on the blower 
shaft is built in one with a friction clutch consisting of four 
bronze sectors which are pressed against the inside of the clutch 
housing by centrifugal force and which come into action when the 
engine speed reaches 600 r.p.m. The casing and its supports, 
partition walls, and diffusors are of aluminum. The super- 
charger complete weighs 105 lb. and was applied to a 260-h.p. 



SUPERCHARGING 



407 



Mercedes engine weighing 925 lb.; it has also been used with 
rotary engines. A four-stage compressor for supercharging up 
to 16,000 ft. weighs 132 lb. 

Brown-Boveri have built a central four-stage supercharger, 
18.5 in. diameter and making 6,000 r.p.m., which gives a periph- 
eral speed of 490 ft. per second. The gear ratio is 4.15 to 1. 
This machine supplies 9,200 lb. of air per hour at 0.52 atmosphere 
initial and 1 atmospheric final pressure; it is driven by a 125-h.p. 
engine and sends compressed air to engines aggregating 1,200 
h.p. The gear teeth are of the Maag form, 0.5 in. circular pitch, 




O 



Fig. 312. — Spring coupling for supercharger drive. 

2 in. face width, are hardened and ground and are loaded to 
1,600 lb. per square inch at full load. Oil is injected directly 
between the teeth. The blower is connected to the engine 
shaft through a leather block joint. The coupling has a disc 
flywheel mounted on it (weight of both 44 lb.) to give smooth 
operation and protect the gears against shock. Tests with 
spring couplings have shown that the actual forces at the gear 
teeth are likely to be four times the normal driving force. The 
details of a successful spring coupling are shown in Fig. 312. 
The efficiencies of these multi-stage compressors (isothermal 
basis) average about 65 per cent; they may reach 68 per cent in 
some cases. 



408 



THE AIRPLANE ENGINE 



With a supercharger direct-coupled to the engine and 
operating all the time, a throttle valve should be placed on 
the suction side of the supercharger — either hand or auto- 
matically operated — to keep down the pressure and power at low 
altitudes. With this location of the throttle valve the power 
absorbed by the compressor will be less than with the throttle 
on the discharge side. With central superchargers, the air 
pressure is controlled by throttling the supercharger engine; 
the air pipe to each engine is fitted with a relief valve, a throttle 



Carburetor 



Automatic Air 
Infer" Va/ve 




Fig. 313. — Explosion relief valve. 



valve for the compressed air and an automatic air admission 
valve which closes when the supercharger comes into action. 
Relief or explosion valves are important; without such a valve 
a back fire would be likely to destroy the partition wall between 
the last two stages of the blower. Figure 313 shows an 
explosion valve held on its seat \>y springs and also an automatic 
inlet valve which closes whenever the compressor is in use. 
The difficulties of geared drives can be eliminated and greater 
total power obtained by the use of an exhaust gas turbine for 
driving the compressor. In this case there is no fixed relation 
between the engine and compressor speeds. The exhaust 



SUPERCHARGING 



409 



manifold leads to the nozzle chamber of the turbine and the 
compressor is mounted on the turbine shaft. The scheme is 
shown in Fig. 314. 

Much development work has been done on turbo-superchargers 



Exhaust io Turbine 




: Turbine Discharge 

T Air Impeller 

Combined Turbine 
"ft and ' Compressor S/rcnf 



Carburetor- 



—\ Carburetor Induction Pipe .. vTy 
V ■*■ -SJ 



Air Compressor 

Inlet- 
Air Discharge to 
Induction Sgstem 



Fig. 314. — Diagram of exhaust- turbine supercharger. 

but they must still be considered as in the experimental stage. 
The principal difficulties encountered have been with the exhaust 
valves, which are subjected to a higher temperature and which are 
not cooled by exposure to the outside temperatures; with the 
manifold, which is kept continuously at a high temperature and 



Turbine Rotor . 




Airlnlet 



Impeller 

.^_^5hafttvr 
Rotating 
Members 

Air Outlet to 
' Carburetor 



Centrifugal 
Comoresser 



Rateau exhaust-turbine supercharger. 



which gives expansion troubles and difficulties in maintaining 
tight joints; and with the nozzle plate and blades of the turbine, 
which are continuously at high temperatures. The increase in 
weight due to the supercharging device for a 400-h.p. engine can 



410 



THE AIRPLANE ENGINE 



be made from 15 to 20 per cent of the engine weight when a 
peripheral speed of 900 ft. per second is used for the compressor. 
The pioneer work on turbo-superchargers has been done by 
Rateau in France. Figure 315 shows a cross-section of his 
arrangement. The results of tests of the Rateau supercharger at 



u 

C 










































C£ 






















£ -2 
2 S 

■£"130 


















































































2£ 






















$* 






















i ?o 























Fig. 316. 



20 30 40 

Pressure in Exhausf Manifold, 
Inches of Mercury , Absolute. 

-Pressures in Rateau exhaust-turbine supercharger. 



an altitude of 9,000 ft. are given in Figs. 316-318. Figure 316 
shows the relation between the back pressure on the engine and 
the pressure at the carburetor; the exhaust pressure stays at 
about 2 in. of mercury above the carburetor pressure. Figure 317 
shows the variation of pressure and temperature ratios in the 



1.6 




























1.5 

o 






















































-+- 












2!S 


p^i 


io^, 












O) 










pT^" 
















3 
if) 

S 1.3 






























s 


+ 




^ 


ffl+i° 
















§ 1.2 






u 


f«fi 




















fei 


«• 




















E 
|2 1.1 






















































10 





























20000 22000 24000 26000 28000 30000 
R.p.m. 
Fig. 317. — Performance of Rateau exhaust-turbine supercharger. 

compressor with varying r.p.m. Figure 318 gives the variation 
of over-all efficiency of the turbo compressor with variation of 
r.p.m.; the turbine efficiency naturally increases as the bucket 
speed approaches the designed speed. Tests of a Lorraine- 
Dietrich 8-cylinder, 160-h.p. engine show an increase of power 



SUPERCHARGING 



411 



from 111 to 164 h. p. at 9,000 ft. altitude by the use of this super- 
charger; the engine speed increased from 1,370 to 1,550 r.p.m. 
Tests of a Breguet plane with a 300-h.p. Renault engine showed 
the time of climb to 16,400 ft. decreased from 47H to 27 min. and 
the horizontal speed at that altitude increased from 91 to 120 
miles per hour by the use of this supercharger. The ceiling was 
increased 13,000 ft. and the speed at the new ceiling was 25 per 
cent greater than that at the old ceiling. 

British tests of a Rateau supercharger fitted to an air-cooled 
engine indicate the possibility of developing within 12 per cent 
of ground power up to a height of 17,000 to 20,000 ft. This is 



0.300 

0.215 

m 0.250 
u 

.f 0.225 
u 

§, 0.200 



0.150 
0.125 

0.100 



































































4 


' 





































































































































































































































































































8000 10000 12000 14000 16000 18000 20000 22000 
Turbine Rpm 
Fig. 318. — Over-all efficiency of Rateau exhaust-turbine supercharger. 



obtained by maintaining ground pressure at the carburetor, 
which was found to entail a back pressure of about 19 lb. abs. 
at the exhaust. 

The Moss turbo-supercharger (General Electric Co.) has been 
designed for, and used successfully on, the Liberty engine. The 
exhaust manifolds, of rectangular form, increase in cross-section 
as they come forward to the front of the engine and join at the 
nozzle box (Fig. 319) which is situated inside the Vee at the level 
of the tops of the cylinders. The nozzles cover about one-half 
the circumference of the wheel. The turbine wheel is 9.1 in. in 
diameter, the compressor 10.5 in.; the peripheral speed is about 
1,000 ft. per second. The turbine and compressor spindle is 
supported at the rear in a water-cooled bearing mounted on the 
intake pipe; at the front,- the bearing is in the air intake to the 
compressor. The compressor is provided with guide vanes 



412 



THE AIRPLANE ENGINE 



(Fig. 319) and discharges at the bottom into the intake pipe 
which extends horizontally backward into the Vee and supports 
the two carburetors. Performance data on this supercharger are 
not available but preliminary tests at Pike's Peak (barometer 
18 in.) showed an increase of engine power from 251 h.p. to 367 
h.p. when running at 1,800 r.p.m. 

The use of a supercharger leads to some complication in the 
fuel supply system. The carburetor float chamber is kept at the 
compressed-air pressure, which is variable, and may be 7 or 8 lb. 




Section A- A Section B~5 




■NojjIeBox 
I ,Water-Cooled Bearing / 



Exhaust Manifold 



\ D 



f — Carburetors 




Air Duct 



'Exhaust Gas Turbine 
'Centrifugal Blower 



Fig. 319. — Moss exhaust- turbine supercharger. 



per square inch in excess of the atmospheric pressure. The fuel 
has to be fed to the carburetor against this pressure. An air 
pressure system is undesirable both because of danger of leakage 
of fuel and also because the tanks would have to be made heavier 
to withstand such high pressures. The pressure in the fuel line 
must not exceed the pressure in the float chamber by an amount 
sufficient to lift the float valves and thereby flood the carburetor; 
the excess of pressure should be less than 5 lb. per square inch. 
A practicable system is to supply the fuel by a direct-acting 



SUPERCHARGING 



413 



,-To Supercharger 
' Air Duct. ■ 



engine-driven fuel pump delivering into a line equipped with a 
spring-loaded relief valve which is subjected to the compressed- 
air pressure on one side and the pump discharge pressure on the 
other. The spring is adjusted to lift at any desired excess of 
fuel over air pressure and by-passes some of the fuel to the suction 
side of the pump. Such a valve is shown in Fig. 320. The air 
is admitted to the inside of a corrugated copper "sylphon;" 
the fuel pressure must be 
sufficient to overcome both 
the air pressure and the spring 
compression. 

Another method is to main- 
tain the supercharger pressure 
in the gravity tank (Fig. 321) 
and to return to the main 
tanks the excess of gasoline 
pumped to the gravity tank 
through a float-operated valve. 

Small fuel tanks with air 
pressure obtained from hand 
pumps are used for starting 
or for emergencies. 

Another method of increas- 
ing the power of an airplane 
engine at high altitudes is by 
supplying the engine with oxy- 
gen. Approximately 4 lb. 

of oxygen is necessary to burn 1 lb. of gasoline. This weight 
is so considerable that oxygen can be carried only for emer- 
gency uses, as, for example, in combat, where it is desired to 
increase the speed of the plane for a short time. As the oxygen- 
gasoline mixture would give excessive temperatures and pressures, 
the oxygen can be used only for enriching the air and permitting 
a moderate increase in the heat developed per cycle. The oxygen 
must be carried in liquid form at atmospheric pressure, otherwise 
the weight of the container becomes excessive. As the boiling 
temperature at ground pressure of liquid oxygen is — 297°F. it is 
necessary to have extraordinarily good heat insulation to keep the 
rate of evaporation down to a permissible figure. A Dewar flask 
is the only practicable insulator in this case, but since it is too 
fragile and brittle to withstand the slapping of the liquid in 




From Pump Discharge 

Fig. 320. — Relief valve for fuel line of 
supercharged engine. 



414 



THE AIRPLANE ENGINE 



flight, it is necessary to put the liquid oxygen in a metal container 
which with proper cushioning is inserted in a Dewar flask. The 
liquid oxygen can be evaporated at any desired rate (1) by 
electrical heating through a coil immersed in the liquid and (2) by 
immersing a metal rod in the liquid to an adjustable depth and 
utilizing the thermal conduction along the rod, or (3) the liquid 
can be siphoned over into the charge going to the cylinder. It is 
evident that special adjustment of the carburetor is necessary to 
meet the condition of oxygen supply. 



Gravity 

Tank-" 



Pressure Compensating Pipe 



Float for 
Operating 
Ya/ves 




Carburetor 



From the 
Supercharger 



ToThe Fuel Tanks 



Fig. 321. — Fuel system for supercharged engine. 



The power developed in a supercharged engine can be absorbed 
satisfactorily only by the use of a variable pitch propeller. A 
four-bladed propeller would be best in order to avoid excessive 
peripheral speeds at high altitudes. The supercharger may be 
used either (a) to maintain constant power, in which case the 
revolutions per minute will vary inversly as the cube root of the 
air density, and the torque will vary as the cube of the density 



SUPERCHARGING 



415 



or (&) to maintain constant torque, in which case the revolutions 
per minute and power will vary inversely as the density. 

A calculation for a machine of standard type with a ground 
speed of 110 miles per hour gives the following results: 





Altitude, 10,000 ft. 


Altitude, 20,000 ft. 




Type of engine 


Maximum 

speed, 

miles per 

hour 


Rate of 
climb, 
feet per 
minute 


Maximum 

speed, 

miles per 

hour 


Rate of 
climb, 
feet per 
minute 


Ceiling, 

feet 




105 
126 
132 


450 
1,050 
1,420 


141 
159 


930 
1,770 


19,000 




61,000 








retical limit 



CHAPTER XVII 



MANIFOLDS AND MUFFLERS 



Air Intakes. — The location, dimensions and orientation of 
the air intake to the carburetor have considerable influence on 
the capacity of the engine. Attempts are often made to increase 
the density of the air going to the carburetor by having the air 
intake face full forward, so as to establish in the intake a pressure 
which is the sum of the surrounding atmospheric pressure and the 
velocity head equivalent to the air velocity of the plane or of 
the slip stream. With a plane making 120 miles per hour at the 
ground this dynamic head would be about 0.24 lb. per square 
inch. The necessity for keeping the air intake pipe of moderate 
length does not usually permit the intake to be located in free air so 
as to take advantage of this velocity head. Inside the fuselage 
there is very small air velocity and the practical means of getting 
access to high velocity air is by extending the air intake pipe 

above the cowling. This arrange- 
ment has the great advantage 
that spillage of gasoline into the 
fuselage is thereby prevented and 
a frequent source of fires is elimi- 
nated. The disadvantage of this 
i i i i i arrangement when applied to the 

\^^S J usual form of carburetor is that the 

air intake pipe will have two or 
more right-angle turns. With an 
inverted type of carburetor, with 
jets discharging downward this dif- 
ficulty is overcome and a short di- 
rect air intake pipe can be employed. 
Investigations in England on the best shape of air intake 
have shown that maximum power is obtained when the intake 
pipe is cut off at an angle of 25 deg. to the horizontal so as to 
form a scoop and that instead of mounting the scoops dead ahead 
they are best turned about 7 deg. to compensate for the propeller 
slip stream (see Fig. 322). The double right-angle turn of the 

416 



'•-CowJing 
[Carburetor 



Fig 




322. — British design 
intake pipe. 



■25 c 



<— ... 

Stand 
Pipe 



^7 



MANIFOLDS AND MUFFLERS 417 

air on its way to the carburetor results in a disturbed air flow 
which can be largely corrected by the insertion of the baffle plate 
indicated. In the Liberty engine a single standpipe has been 
used to serve two carburetors, by the use of the duplex air intake 
shown in Fig. 323. 

The length of intake pipe will affect the engine capacity 
because the inertia of the air column tends to make the flow to the 
carburetor continuous in spite of 
the intermittent action of the 
cylinder suctions. In the Lib- 
erty engine the optimum length 
with inverted type of carburetor 
is found to be 6 in. at normal 
revolutions per minute. 

Intake Manifolds. — The main 
function of the intake manifold 
is the distribution of the mixture 
formed in a carburetor to several FlG - 223 '" Uh ^ ff ne duplex in " 
cylinders. For good efficiency 

and capacity the strength and density of the mixture reaching 
all cylinders should be the same. To ensure these results the 
amount of fuel entering, and the pressure drop from carburetor 
to inlet valve, should be the same in all branches. With a fuel 
which is not completely vaporized before the branching of the 
manifold occurs it is very difficult, if not impossible, to ensure 
proper distribution of the fuel. This difficulty increases as the 
volatility of the fuel decreases and is therefore greater with 
ordinary commercial gasoline than with airplane fuels. 

For the vaporization of a volatile fuel the important factors 
are (1) air supply of sufficiently high temperature, (2) fine 
atomization at the jet, (3) avoidance of obstacles in the path of 
the mixture so that the atomized liquid drops may have no 
opportunity for coalescence before being vaporized, and (4) 
sufficient time for the vaporization. The temperature of the 
mixture after vaporization is much below that of the entering 
air because the latent heat of vaporization of the fuel is taken 
from the air. Assuming a latent heat of 135 B.t.u. per pound, 
complete vaporization will produce a temperature drop of 
47°F. with an air fuel ratio of 10 to 1, and a drop of 25°F. with 
an air fuel ratio of 20 to 1. If all the fuel is not vaporized the 
temperature drop will be less. The fall of temperature in the 

27 



418 



THE AIRPLANE ENGINE 



manifold may bring about a deposit of ice both inside and 
outside the manifold as a result of the freezing of the moisture 
in the air. It has been suggested 1 that accumulation of ice 
inside the manifold may be the cause of numerous unexplained 
engine failures and resulting crashes. The relation between the 
manifold temperature, measured at the intake valve, and the 
air supply temperature in a Liberty engine is shown in Fig. 324. 
The manifold was water-jacketed. Below 20°F. there was so 
little evaporation that the manifold temperature was higher 
than the air supply temperature. As the air supply temperature 
increased up to 120°F. the temperature fall increased up to 
about 35°F. The air fuel ratio was about 16 to 1. 



iH 

Q 



35 

o 



120 

100 
80 
60 
40 
20 




• J 

~X~jr 

£ 



Fig. 324. 



20 40 60 80 100 120 
Air Intake Temperature,DeqFahr. 
■Temperature change in intake manifold. 



With air initially cold it will not be possible to vaporize the 
fuel completely by heat absorbed from the air because at low 
temperatures the air will become saturated with the fuel vapor 
before all the fuel is vaporized. Table 14, page 229, shows, for 
example, that with a theoretically correct mixture, the air cannot 
hold all the pentane in the vapor form at a temperature below 
about 38°F. In such case the only chance for complete vapor- 
ization is by supplying heat from outside. This is accomplished 
by utilizing some of the heat either of the jacket water or of the 
exhaust gases. The possibilities are (1) to preheat the air before 
it enters the carburetor, (2) to heat the mixture in the manifold, 
and (3) to heat the manifold locally at some place on which the 
1 Sparrow, Technical Note, No. 55, Nat. Adv. Co-nun. Aeronautics. 



MANIFOLDS AND MUFFLERS 



419 



liquid drops impinge so as to supply heat to the liquid only 
(hot-spot method). 

All preheating is objectionable in that it diminishes the density 
of the charge and thereby decreases the capacity of the engine. 
That method of preheating is best which causes vaporization 
with the minimum resulting temperature of the mixture. All 
three methods of preheating are employed in airplane practice. 
In some of the German engines preheating the air is accomplished 
by taking the air through pipes in the crankcase (Fig. 78), which 
has the advantage of cooling the lubricating oil. The more 
common procedure is heating the manifold by jacket water. 





c^^ri 




Fig. 325. — Liberty engine intake manifold. 

There is no general consensus of opinion as to the best form of 
manifold. It is desirable that sharp turns should be avoided 
as far as possible, that the various branches should have approxi- 
mately the same length, that sudden enlargements should be 
avoided and that the velocities should be high enough to prevent 
deposition of liquid drops but not so high as to cause a large 
frictional resistance. Mean velocities of 120 to 200 ft. per minute 
are common but values up to 250 ft. per minute have been used 
successfully. 

Manifolds usually divide themselves into two classes, the 
short-branch type and the long-branch type. The standard 
Liberty manifold (Fig. 325) is a good example of the short-branch 



420 



THE AIRPLANE ENGINE 



type; it is water-jacketed and has a baffle plate opposite to the 
inlet to equalize the lengths of the three branches. The Benz 
manifold (Fig. 326) is an example of an unjacketed long-branch 
manifold with all three branches of the same length and with long- 
turn elbows. Another design for accomplishing the same purpose 
is shown in the Hall-Scott engine (Fig. 63). 

Manifolds for vertical engines can usually be arranged in 
any way the designer likes. In the Maybach engine (Fig. 80) 
the carburetors are at the ends and the manifolds run along the 
side of the engine with no attempt to equalize the lengths of the 
branches. Ordinarily such arrangements as those of Figs. 325 and 
326 are used. 

In 90-deg. Vee engines there is plenty of room in the Vee to 
accommodate the carburetors and the intake is usually inside the 




Fig. 326. — Benz intake manifold. 

Vee. With this location a short-branch manifold must be used. 
The Hispano-Suiza engine (Fig. 50) shows a typical arrange- 
ment with the transverse pipe water-jacketed. If long-branch 
manifolds are to be used they must either be placed in the rear of 
the engine,. as in the Curtiss engines, Figs. 59 and 62, or the 
intake valves must be on the outside of the Vee. 

In 60-deg. and 45-deg. Vee engines the space inside the Vee is 
small and a favorable design of manifold is difficult if the car- 
buretors are placed inside the Vee. If the inlet valves are placed 
outside the Vee, the exhaust pipes will be crowded inside the Vee 
and may give rise to troubles caused by their proximity to the 
valve springs, etc. An alternative arrangement is to provide a 
space between the two central cylinders of each block (as in the 
Bugatti (Fig. 67) and Fiat engines (Fig. 76)) and to lead the 
induction pipes from carburetors mounted outside the Vee 
through these spaces to manifolds inside the Vee. 



MANIFOLDS AND MUFFLERS 421 

In radial and rotary engines the distribution problem is com- 
paratively simple, especially where an induction chamber is 
provided in the crankcase. The special distributing chamber of 
the Bristol " Jupiter'' engine (Fig. 148) is noteworthy. 

Exhaust Manifolds. — The function of the exhaust manifold is, 
primarily, to conduct the exhaust gases away from the airplane 
with the minimum back pressure at the engine and without fire 
risk to the airplane or annoyance from the discharged gases to 
the pilot. An additional function may be to muffle the sound of 
the exhaust, although this has usually been considered unimpor- 
tant in military machines. The manifold is usually required to 
have a clearance of 2J^ in- from wooden parts and of 3J-^ in. from 
fabric parts of the airplane. 




Fig. 327. — Hall-Scott exhaust manifold. 

The simplest exhaust piping is either its complete absence 
as in rotary engines and certain stationary engines or the use of 
short tubes discharging outwards or upwards. In some engines 
these stub tubes are cut off on the outer end at an angle of 45 
deg. so as to discharge backwards as well as outwards. The 
absence of exhaust pipes or the use of short straight pipes is 
advantageous not only in reducing back pressure but also in 
permitting better cooling of the exhaust valves by radiation and 
avoiding heating of the exhaust valve springs. This arrangement 
does not conduct the gases away from the crew of the airplane. 

A method of discharging the gases overhead and to the rear 
with small back pressure is shown in the Hall-Scott manifold of 
Fig. 327. This arrangement obstructs the view of the piloC in a 
tractor machine but diminishes the noise heard from below. 
Long radius curves are very essential for all the branches if back 



422 



THE AIRPLANE ENGINE 



pressure is to be kept down; the radius should be about 2^ times 
the diameter of the pipe. The arrangement of Fig. 328 with 
discharge to the rear and with an exhaust main of increasing 
diameter not only carries the gases away from the crew but slows 
down the gases before exit and thereby tends to diminish noise. 
As the exhaust period lasts more than two-thirds of a revolution 
there is overlapping of exhausts in a manifold connecting three or 
more cylinders. The velocity of the gases immediately after the 
opening of the exhaust valve is extremely high and its effect on 
the exhaust from any other cylinder whose exhaust valve is open 
at that time should be carefully considered. By the use of two 
concentric exhaust mains to which the cylinders are connected so 




Fig. 328. — Hall-Scott exhaust manifold. 



that no two consecutive exhausts go into the same main, an 
ejector effect can be obtained from the action of a newly opened 
exhaust on the exhaust which is closing, which may cause substan- 
tial scavenging in the closing cylinder. 

The cross-section area of manifold branches is governed by the 
size of the ports in the cylinders; an average value is about 0.15 
sq. in. per brake horse power of the cylinder. 

Mufflers, to be efficient, must slow down the exhaust gases to 
velocities below that of sound (1,100 ft. per second); actual gas 
velocities probably exceed 2,000 ft. per second. For airplane use 
the muffler must be of light weight and yet durable. The con- 
structions employing reversal of direction of the gases and dis- 
charge through small holes are likely to give excessive back 
pressure. Tests by Diederichs and Upton show that the volume 
of the muffler should be about three times that of a single cylinder 



MANIFOLDS AND MUFFLERS 



423 



of the engine, and that the inlet to the muffler should be tangen- 
tial so as to give the gas a whirling motion. For durability the 
muffler should be attached to the end of a tail pipe 6 or 8 ft. long 
which will cool the gases sufficiently to prevent excessive oxida- 
tion of the muffler. One of the simplest and most successful 
mufflers is shown in Fig. 329. The tangential inlet pipe starts of 
circular cross-section and flattens out in a fan shape so as to give 
admission for almost the whole length of the muffler. An inner 



In lei 



Inlet 




Fig. 329. — Diederichs and Upton exhaust muffler. 

shell, AB, is provided with a large number of holes except in the 
small arc between A and B; these holes are smallest near B and 
increase in diameter from B to A (contra-clockwise). The gas 
passes through these holes and then through more holes in the 
innermost shell and finally escapes at the open end of the inner- 
most shell. With this muffler the exhaust noise can be reduced 
80 per cent with about two-thirds of 1 per cent reduction in 
engine power. Mufflers of this type are found to give less back 
pressure than those with axial admission. 



CHAPTER XVIII 



STARTING 



The starting of an airplane engine depends on three things, 
(1) obtaining an explosive mixture in the cylinder., (2) a device 
for igniting it, and (3) a device for turning the engine over. 

Hand Starting. — The idling device on the carburetor will fur- 
nish a rich mixture to the cylinders if the throttle valve is closed 
and the engine is turned at a sufficient speed. If the revolutions 
per minute of the engine is too low (below 20 to 30 r.p.m.) the 
velocity of the air in the intake manifold will be insufficient to 
carry the fuel into the cylinders; this will be the usual condition 




Fig. 330. — Priming system for a 12-cylinder Vee engine. 



with hand starting. To overcome this difficulty the cylinders may 
be primed through individual priming cocks or through such a 
priming system as is shown in Fig. 330. With the engine cold 
only a small part of the gasoline will be vaporized and the rest 
will go as liquid into the cylinder and will dilute the lubricant. 
In very cold weather there will be difficulty in vaporizing enough 
of the gasoline to form an explosive mixture unless the jacket is 
filled with hot water or some other device has been used to heat 
the engine. In such a case a number of attempts may have to be 

424 



STARTING 425 

made before the engine starts, and as an excess of gasoline is put 
in the cylinder before each attempt the walls may be washed 
clear of lubricant and scoring of the cylinder may occur when 
the engine finally starts. 

To overcome these difficulties a more volatile fuel such as 
ether may be used for priming. A more satisfactory practice, 
for cold weather, is to use hydrogen as the starting fuel. This 
has the great advantage that it does not have to be vaporized and 
that it forms an explosive mixture throughout a great range of 
strengths. A hydrogen-air mixture will explode if the hydrogen 
forms from 10 to 66 per cent of the mixture by volume; with 
gasoline vapor the range is only from 1.5 to 4.8 per cent. The 
hydrogen may be admitted through the priming system as 
shown in Fig. 330; it is best to take the gas from a fabric balloon 
in which it is at atmospheric pressure and not from a high-pres- 
sure bottle which would be likely to give an excess of gas. The 
hydrogen must be shut off shortly after the starting as it gives 
more violent explosions than gasoline. 

The regular magneto will not turn fast enough with hand 
starting to give a spark. If a battery system of ignition is used 
there is no difficulty from this source. If a magneto system is 
used it is customary to supply a hand-operated starting magneto 
which is geared to run at high speed and is turned independently 
of the engine. With hand starting the engine may be pulled 
over several times to fill the cylinders with explosive mixture, 
after which a shower of sparks is sent from the starting magneto 
through the distributor to the fully retarded cylinder which is 
ready to fire, or the ignition may be left on, fully retarded, while 
the propeller is pulled over either by hand or by rope. 

The directions for starting the Liberty engine by hand are as 
follows: Inject )4 oz. of lubricating oil through each priming 
cock. Turn switch "off." Turn the engine over five times. 
Open throttle slightly. Retard spark fully. Prime each cylinder 
twice. Turn engine over twice. Turn on one switch. Pull 
down and forward on propeller blade. After the engine starts: 
Advance spark half way. Turn on both switches. Leave 
throttle undisturbed for 5 min. The lubricating oil should be 
warm by the time the jacket outlet water has reached 150°. 
If, in cold weather, it is not, stop the engine for 5 min. ; then start 
all over again. Accelerate and slow down the engine a few 
times. Note that the oil gage registers pressure, 5 lb. at 600 



426 



THE AIRPLANE ENGINE 



r.p.m. At this speed, the ammeter should show " discharge." 
At 1,000 r.p.m. it should indicate " charge." After 5 min. more, 
open the throttle wide. The speed should rise to about 1,600 
r.p.m. 

The necessity of turning the engine over preliminary to 
starting can be avoided by the use of a mechanism which will 
lift the valves and permit the pumping of an explosive charge 
into the cylinders. In the Maybach engine this is accomplished 

(Fig. 331) by lifting aU the 
tappets (inlet and exhaust) 
off their cams by depressing 
the hand lever, A, which 
rotates the two lay shafts, 
BB, and lifts the tappets 
through slots in the lay 
shafts engaging with small 
lugs projecting from the tap- 
pets. At the same time the 
shutter, C, in the exhaust 
main is closed. The hand 
pump, E, is then operated 
and air is sucked through the 
carburetor and the intake 
manifold and through the 
cylinder to the exhaust man- 
ifold and to the pump. When 
the cylinders are thus charged 
the lever, A, is brought ver- 
tical, which restores the engine to its operating position. The 
starting magneto is then used to start the engine. The whole 
operation can be carried out from the pilot's seat. 

The starting torque in high-power airplane engines is very 
considerable. The load consists of two parts, (1) the compression 
load, and (2) the friction load. The compression load increases 
with the cylinder diameter and the compression ratio. The 
maximum torque does not change with the number of cylinders 
because high-compression pressure will exist at any moment in 
one cylinder only. As the compressed gas re-expands the mean 
torque for two revolutions does not increase with increase in 
number of cylinders. Friction is the principal resistance in 
starting the engine, especially in cold weather when the lubricating 




Fig. 331.- 



-Starting mechanism of May- 
bach engine. 



STARTING 



427 



oil may be near the solidifying point. Tests at McCook Field 1 

to determine the average starting torque for various engines have 

yielded the following results. The average air temperature was 

Average Starting Torque, Pounds-feet 





02 

f-l 
0) 

.5 
% 

u 


& 

o 

ft 

Pi 

o 

73 
<u 


a 

09 

s 


tn 

a; 

o 

p. 
o 
pq 


to 

<D 

o 

a 

M 

o 

Pi 
+3 

m 


DO 

a> 
fi 
'3b 
a 

<u 

"3 

Pi 

a is 


CO 

1? 

Pi 
0> 


Starting torque, 
pounds-feet 


Engine 


Throttle 
open 


Throttle 
closed 




■a a 


& Si 








12 

8 
6 
8 
6 


400 
300 
210 
180 
185 


1,700 
1,800 
1,700 
1,600 
1,400 


5.00 
5.51 
5.00 
4.75 
5.90 


7.00 
5.91 
7.00 
5.25 
7.08 


2 
2 
2 
1 
1 


6 
6 
6 
3 
3 


130 
106 
133 
110 
150 


124 
102 
105 
82 
139 


156 
101 
135 

87 
153 


143 


Hispano-Suiza 


96 
110 




77 


B. M. W 


145 







75°F. With freezing temperatures the results would have been 
much higher — probably doubled. It should be noted that the 
starting torque does not increase nearly as rapidly as the engine 
horse power. 




Fig. 332. — Compression release of Benz engine. 

From the preceding table it may be presumed that with low 
temperature the mean starting torque for a 400-h.p. engine will 
be from 300 to 400 lb. -ft. It is difficult for a mechanic to exert 
this torque on the propeller even in a land plane without danger to 
himself from overbalancing; in a sea plane it is even more difficult 
to accomplish. The starting torque can be diminished by reduc- 
ing or eliminating the compression. For this purpose compres- 
sion release cams may be provided on the camshaft to keep the 
exhaust valves open during the compression stroke. In the 
Benz engine (Fig. 332) the compression release cams are brought 

1 Air Service Information Circular, No. 126. 



428 



THE AIRPLANE ENGINE 




Fig. 333. 



-Compression release of Basse- 
Selve engine. 



into action by the axial movement of the camshaft effected by a 
square-thread screw operated by a small lever at the rear of the 
crankcase. The camshaft is returned to its normal running 
position by a spring inside the front end of the shaft. The relief 
cams open the exhaust valves 35 deg. early and close them at 
22 deg. late. 

In the Basse-Selve engine (Fig. 333) the compression release is 

operated by means of a rod 
which lies horizontally along 
the outside of the camshaft 
casing directly underneath 
the exhaust-valve rocker 
arms. The rod is slotted in 
such a way (see separate 
detail, Fig. 333) as to form 
cams which lift the exhaust 
valve rockers when the rod 
is partially rotated by means 
of the hand lever at the end of the rod. With this device the 
exhaust valves remain open so long as the rod is in the rotated 
position. 

In addition to the difficulty which is experienced in starting 
large engines by the propeller there is a considerable element of 
danger, which has proved fatal in many cases. To obviate this, 
recourse may be had to a portable engine cranker, which is usually 
available only at airdromes, or to a starting mechanism integral 
with the engine. Whatever the nature of the starting mechanism 
it should be thrown automatically out of action as soon as the 
engine fires and it should also go out of action if the engine fires 
before the dead center and starts to turn backward. The con- 
nection from the starter to the engine is 
through a dog or clutch, as in Fig. 334, 
which is pushed out of mesh when the 
engine starts forward but will remain in 
mesh if the engine starts backward. FlG - 334.-Dog clutch. 

The acceleration of the engine shaft as a result of firing one 
of the cylinders is very great and if the engine starts backward 
this acceleration will be transmitted to the starting mechan- 
ism. As the starting mechanism is always geared, with a ratio 
which may be 100:1 or more, the motivating element will be given 
an enormous acceleration which will produce high teeth pressures 




STARTING 



429 




m\ 



Fig. 335. — Safety clutch. 



in the gears and will probably strip the teeth of the gears or the 
dog projections of the clutch. If the gears are hand-operated 
through a crank the gear ratio may be 10 or 20 to 1, and, although 
the teeth may hold on a back fire, the mechanic will be endan- 
gered by the high speed of the handle. To safeguard the starting 
gear some kind of friction clutch or other safety drive should be 
placed close to the dog. Multiple-disc clutches are suitable on 
account of their compactness and low weight, but all such friction 
devices are uncertain in their action. 

A different type of safety device with an oscillating member 
is shown in Fig. 335. l The oscillating 
member, A, takes no part in the trans- 
mission of the load. The engine drive 
is to the right of the figure. In normal 
rotation, A is driven by the gear wheel, 
C, and is free to oscillate as determined 
b} r the tapered sides of the stationary 
projections on D. If a reverse rotation 
occurs D prevents the rotation of A, which causes the sleeve 
on which C is mounted to move to the right, and thereby throws 
C out of mesh. As the action cannot take place until some back- 
ward movement has occurred this gear should be placed on a 
shaft geared to the engine shaft, so that release may occur 
quickly. In a hand-operated system it might be on the handle 
shaft. 

Integral cranking mechanisms are either operated by hand, 
by compressed air, or by electric motor. It is usually necessary 
to have the engine rotating at a speed of from 10 to 20 r.p.m. 
before regular operation will take place. Hand mechanism must 
have the operating handle at the side or rear of the engine and 
consequently must employ worm, bevel, or helical gears, usually 
in conjunction with spur gears. The efficiency of the gear train 
is poor and the attainable cranking speed very low. The hand 
mechanism for the Hispano-Suiza engine is shown in Fig. 336. 
It includes double-reduction spur gears, a dog clutch, a releasing 
spring and a gear-driven starting magneto. 

Compressed air may be utilized (1) in a motor which cranks 
the engine, (2) it may be carbureted and sent through a dis- 
tributor and exploded in the cylinders, or (3) it may go direct 
to the cylinders at a high pressure through a distributor and 

1 Sherman, The Automobile Engineer, December, 1919. 



430 



THE AIRPLANE ENGINE 



special non-return valves. If a compressed air motor is used, it 
can be run from the engine as an air compressor to store up the 

^&~- Starting 

n Magneto 




Fig. 336. — Hand-starting mechanism for Hispano-Suiza engine. 




<- Engine Face 



Fig. 337. — Radial air compressor. 




air necessary for starting. In the Motor-Compressor Company's 
starter (Fig. 337) a multi-cylinder radial compressor is lined up 



STARTING 



431 



with the crankshaft at the rear of the engine. When operated 
as a compressor it is directly connected through a positive clutch 
to the engine and runs at engine speed, compressing air up to 
230 lb. pressure in a wire-wound tank. When used as a starter 
it drives the engine through a train of spur gears and rotates at 
seven times engine speed. Automatic arrangements are provided 
for throwing out the compressor when the air pressure reaches 
230 lb. and for throwing out the cranker when the engine starts. 
The apparatus weighs about 50 lb. for a 200-h.p. engine. 




Fig. 338. — Compressed air distributor. 



In the Christensen sj T stem compressed air is sent through a 
special carburetor, a distributor and non-return valves to the 
cylinders. The air pressure is sufficient to start the engine 
(say 100 lb.) and a retarded spark gives a late explosion and 
initiates regular operation. This system uses a minimum of 
compressed air for the starting. An air compressor driven by a 
hand-operated clutch from the crankshaft and an air tank are 
necessary parts of the system. The whole weighs about 40 lb. 
for a six-cylinder engine. 



432 



THE AIRPLANE ENGINE 



Compressed air can be used inside the cylinders without 
carbureting. If no air compressor is provided a steel air tank 
must be carried of sufficient capacity for several starts. This 
arrangement uses a distributor and non-return valves; it will 
weigh less than the Christensen system, but is likely to leave the 
aviator stranded on occasion. 

The details of an air distributor are shown in Fig. 338. When 
the starting lever, H, is thrown in, the central main air valve, 
B } is opened by the boss at the end of the sliding sleeve and the 
individual valves, D, E, admitting air to the individual cylinders 
are opened in turn by the rotation of the face cam, G. 

Another method which has been used but is now abandoned 
is to insert a black powder cartridge in a special fitting in the 
cylinder head. On detonation this will give a considerable 
pressure and should start the engine. It causes a deposit of 
carbon in the cylinder and is not adapted to remote control. 

Electric cranking has been used considerably. It requires a 
12-volt battery which will normally have to supply 100 amperes 
but may be called upon to give to 200 amperes for a half minute 
or more; an electric motor which will operate at 3,000 to 4,000 

r.p.m., carrying a small 
spur gear on the armature 
shaft; and a double reduc- 
tion gear with a speed 
reduction ratio of 100 to 
150. The last gear is pref- 
erably fastened to the rear 
propeller hub flange, the 
motor reduction gears 
being mounted on the 
crankcase between the 
front cylinders and the propeller hub. The gears can be brought 
into mesh by the use of a solenoid wired in series with the motor. 
The Bijur starter used on Liberty-12 engines has a six-cell battery 
weighing 35 lb. with a capacity of starting the engine 150 times 
under normal conditions. The starting motor and gear may 
weigh 24 lb. and give an engine speed of 40 to 50 r.p.m. The 
maximum torque available at the engine crankshaft is 1,300 
lb. -ft. An arrangement in which the gears are kept inside the 
nose of the engine is shown in Fig. 339 , a this arrangement sup- 
1 Sherman, loc. cit. 




V//////A 



Fig. 339. — Electric starter. 



STARTING 433 

poses the starter incorporated in the engine design and not 
adapted, as usual, as an afterthought. If this starter is mounted 
at the rear of the engine it can be made a combined hand and 
electric starter by putting a dog clutch and cranking handle on 
the intermediate shaft. 

Various portable crankers have been devised for airdrome use. 
The U. S. Air Service has used an electric cranker driven by an 
automobile-starting motor and storage battery through double- 
reduction gearing, and mounted on a motor truck. As the weight 
of the gearing does not have to be considered in this case it 
has been found unnecessary to provide an automatic release in 
case of engine back fire. The cranker is mounted in a spherical 
bowl permitting universal adjustment, it is brought up to the 
end of the propeller shaft and adjusted so that its shaft is in line 
with the engine crankshaft. An engagement lever then pushes 
the perforated face plate at the end of the cranker shaft against 
the front propeller hub flange when some of the nuts enter the 
perforations. The engagement lever is withdrawn as soon as 
the engine starts. 

The Odier portable starter, which has had considerable use, 
employs a long single-acting steel cylinder and piston operated 
by carbon dioxide from a steel bottle of liquefied carbon dioxide. 
The piston carries a pulley at its free end, over which a cable is 
passed with one end fastened to the cylinder and the other 
wrapped around a drum and then fastened to an elastic cord. 
The drum has four bolts placed symmetrically around the 
peripherjr at one end — parallel to the drum axis — projecting 
sufficiently to engage a kind of dog clutch on the front propeller 
hub flange. The cylinder is carried on an inclined wooden 
arm and a vertical leg of adjustable height, so that the bolts can 
be brought up to the propeller hub level. The high pressure of 
saturated vapor of carbon dioxide (308 lb. per square inch at 
0°F.) provides a large starting force in a cylinder of small diam- 
eter. The piston stroke is such as to give two revolutions of 
the engine with high speed. The cranker is thrown forward and 
out of mesh when the engine starts by the action of the dog teeth 
on the propeller flange. In case of a back fire the piston is 
pulled back, the gas is recompressed, and the elastic cord becomes 
slack and permits the drum to revolve freely. The weight of the 
whole apparatus is 44 lb.^o that it can be carried in the plane if 
desired. 

28 



CHAPTER XIX 

POTENTIAL DEVELOPMENTS 

Increased Compression. — The best airplane engines give a 
notably better performance, both as regards fuel consumption 
and horse power per unit of piston displacement, than other 
gasoline engines. The possiblity of this improved performance 
results from the use of a higher compression ratio, which in turn 
is only possible through the use of the volatile aviation gasoline. 
Ordinary commercial gasoline containing larger fractions of the 
heavier paraffines (nonane and decane) would detonate at the 
compression pressures reached in airplane engines. The de- 
pendence of fuel consumption on the compression ratio is shown 
in the following table, 1 which gives the theoretical consumption of 
gasoline (lower heat value 18,600 B.t.u. per pound) per brake 
horse-power hour with an assumed mechanical efficiency of 90 
per cent and with variable specific heats. 



Ratio of compression 

Gasoline per brake horse-power 
hour 



4.0 



0.416 



4.5 



0.398 



5.0 



0.375 



5.5 
0.361 



6.0 
0.350 



The best modern engines use 0.45 to 0.50 lb. per brake horse- 
power hour with a compression ratio around 5.0; that is, the 
attained efficiency is 0.375 ■*■ 0.45 = 83 per cent of the theore- 
tical efficiency. With higher compressions this relative efficiency 
tends to increase. At maximum load, which is obtained only 
with richer mixtures (see p. 33), the relative efficiency averages 
75 per cent for water-cooled and 76.5 per cent for radial air-cooled 
engines with aluminum cylinders. 

The use of still higher compression pressures, and consequently 
higher efficiencies, is possible by the use of gasoline mixed with 
toluol, alcohol or other substance which will prevent detonation 
(see p. 236). This field of improvement is now under active 
investigation and promises considerable improvement. 

1 Gibson, Trans. Royal Aeronautical Society, 1920. 

434 



POTENTIAL DEVELOPMENTS 



435 



Use of Inert Gases. — A high compression pressure can be used 
without danger of detonations, and consequent preignitions, by 
taking in cooled exhaust gases with the charge. The influence 
of such admixture is shown in Fig. 340, which is taken from 
Ricardo's tests. 1 With a high-grade fuel which, when operating 
at full throttle and with an economical setting, detonates at a 
compression ratio of 4.85:1, the full power can be maintained by 
admitting inert gases in sufficient quantity to prevent detonation 
up to a ratio of compression of 6:1. That is, the decrease in 
weight of fresh charge taken in is fully compensated by the in- 
crease in engine efficiency up to that ratio of compression. If 
still higher compression is used (for an oversized engine for high- 

150 

CL 
LU 

^140 

-15 

-£ 

J 130 

120 
i 0.6 
1.0.5 

0.3 

4.0 4.5 5.0 5.5 60 6.5 7.0 7.5 8.0 
Compression Rcch'o. 

Fig. 340. — Effect of addition of inert gases on engine performance. 

altitude flight, see p. 390) the power will fall off* The increase 
in economy by the use of the inert gases, with increase of com- 
pression ratio from 4.85 to 6, is seen to be about 6 per cent. The 
dotted lines show the performance obtained using a fuel doped to 
prevent detonation and without the admission of inert gases. 
With a compression ratio of 7:1 the exhaust-controlled engine 
develops 84 per cent of the power which would be developed by a 
pure non-detonating mixture. 

Any increase in efficiency from increase of compression ratio 
will also increase the power output in practically the same ratio 
and is therefore doubly valuable. 

1 Trans. Royal Aeronautical Society, 1920. 















__^- 












«» 


*•*" 










^ 


_ZH 


e.p 




























































L^ 


tiCor_ 


isurrv^ 


fion^. 







































436 



THE AIRPLANE ENGINE 



The maximum brake m.e.p. possible for an engine with inlet 
valve closure of 50 deg. late, with volumetric efficiency of 88 per 
cent and mechanical efficiency of 90 per cent, is as follows: 1 



Nominal compression ratio . 
Maximum brake m.e.p 




The best recorded results for both air- and water-cooled engines 
with compression ratios of 4.5 to 5.0 are very close to these 
figures. 

Tests of single engines have shown consistently better results 
than those of multi-cylinder engines. The best air-cooled single- 
cylinder engines have shown a relative thermal efficiency of 91 
per cent. The difference between the best performance of a 
single-cylinder water-cooled engine and the performance of a 
12-cylinder Vee engine is from 8 to 10 per cent. The difference 
depends on the efficiency of the induction system and represents 
the possible saving by better distribution in the induction system. 

As the efficiencies of the best single-cylinder engines using 
weak mixtures are within 10 to 15 per cent of the theoretical 
maximum, it is evident that little further progress is possible in 
improving the thermal efficiency of engines using the present 
cycle of operations. Increase in capacity (b.h.p. per unit of 
piston displacement) can be obtained by supercharging or by im- 
proving the volumetric efficiency. This latter method offers some 
chance for improvement as the measured values vary from 70 to 
85 per cent with exceptional values up to 90 per cent. 

The fire risk in airplanes could be practically eliminated by 
the use of kerosene as fuel. Kerosene may be used either (1) by 
vaporizing it outside the engine, or (2) injecting it into the cylin- 
der as a liquid either during the suction stroke or at the end of 
compression. To vaporize a reasonable proportion the initial 
temperature must be not less than 140°F., which results in a 
reduction in the weight of the charge of about 20 per cent as 
compared with gasoline and a corresponding decrease in engine 
power. Furthermore, it is chemically much less stable than gaso- 
line and detonates at a lower temperature so that a lower com- 
pression ratio must be used, which further diminishes the power 
and decreases the efficiency. The heavier fractions condense 
on the cylinder wall and, passing into the crankcase, thin the 

1 Gibson, loc. cit. 



POTENTIAL DEVELOPMENTS 437 

lubricating oil. Injection of the fuel during the suction stroke 
intensifies this last trouble but reduces the loss due to preheating. 
Injection at the end of compression presents many difficulties 
common to Diesel engines, which are not yet satisfactorily solved. 

Modifications of the Otto cycle would seem to offer considerable 
possibilities for increased efficiency. The pressure at the opening 
of the exhaust valve is usually from 60 to 70 lb. per square inch. 
If more complete expansion could be obtained a considerable 
increase in efficiency might be effected. Attempts have been 
made to realize this potential increase in work along two lines, (1) 
by more complete expansion in the cylinder and (2) by expanding 
the gases after leaving the cylinder. 

1. A lower terminal pressure in the cylinder can be obtained 
either (a) by throttling or cutting off the admission of the charge 
or (6) by making the expansion stroke longer than the com- 
pression stroke. The former method (a) is that of the oversized 
engine (p. 390) and is employed primarily for maintaining power 
at high altitudes. Its use requires a larger and therefore heavier 
engine for a given capacity, which is a serious detriment. The 
method (b) can be carried out by the use of a variable -stroke 
engine and has the incidental advantage of permitting a more 
complete scavenging of burned gases. 

The Zeitlin engine which is now under development is an 
example of a variable stroke engine, but this feature is utilized in 
this case only to give better scavenging and thereby to permit the 
admission of a greater weight of charge. It is a single-valve, 
nine-cylinder air-cooled rotary which follows the Gnome engine in 
taking in its air supply through the open exhaust valve and 
mixing with it an overrich mixture through ports uncovered by 
the piston near the end of the suction stroke. The variable 
stroke is obtained by mounting, on the crankpin, eccentrics, which 
are driven by gearing around the crankpin in the same direction 
as the engine but at one-half engine speed. The connecting rods 
are mounted on these eccentrics and consequently the piston 
travel will vary throughout two revolutions of the engine. In the 
engine with 107.75 mm. crank throw, the working stroke of the 
engine is 181 mm. ; the exhaust or scavenging stroke is 203.5 mm. ; 
the suction stroke is 226 mm. and the compression stroke 203.5 
mm. As the admission ports are not covered until the piston has 
made part of the compression stroke the effective compression 
stroke is practically the same as the working stroke. The admis- 



438 THE AIRPLANE ENGINE 

sion ports in the cylinder are uncovered near the end of the suc- 
tion stroke. The length of the scavenging stroke is such as to 
clear the cylinder almost completely of burned gases, so that the 
new charge is undiluted by them. 

This engine is arranged to give variable compression by open- 
ing the exhaust valve during the early portion of the compression 
stroke and permitting some air to escape before it has time to 
mix with the overrich mixture. It is designed for a maximum 
compression ratio of 7 and has controls for decreasing this to 
4.5. The maximum compression is for altitudes of 10,000 ft. or 
more; the minimum compression is for ground operation. The 
strength of the mixture will obviously change with the ratio of 
compression; the mixture will be rich at the ground and will be 
leaned to maximum economy strength at normal flying level. 

2. The further expansion of the gases after leaving the cylinder 
can be carried out either in a gas turbine or in a reciprocating 
engine. The use of the exhaust gas turbine for driving a super- 
charging blower has already been discussed (p. 408) ; in view of 
the high speed of the turbine shaft, which is necessary if the 
turbine is to have fair efficiency, it is doubtful whether such a 
turbine could be geared down so as to help drive the propeller 
without excessive loss of power. For driving high-speed auxil- 
iaries such as the supercharging blower it has. an important 
field. 

Compounding. — Expansion of the gases in a reciprocating 
engine can be accomplished by following the methods used in 
compound steam engines. The problem is simplified in some 
respects in the gasoline engine because one double-acting low- 
pressure cylinder can take the exhaust from four high-pressure 
cylinders and as the temperature of the gases is greatly reduced 
by the time they reach the low-pressure cylinder it might be 
possible to operate without trouble from excessive piston tem- 
perature. As the friction work in the high-pressure cylinders is 
equal to about 15 lb. per square inch of piston area there should 
be a pressure drop of at least that amount at exhaust, or with 
70 lb. terminal pressure in the high-pressure cylinder the receiver 
pressure should be about 50 lb. Furthermore, in order to get a 
full charge in the high-pressure cylinder it would be necessary 
to have an atmospheric exhaust from the high-pressure cjdinder 
at the end of the exhaust stroke and immediately after closure 
of the exhaust to the receiver. The piston displacement of the 



POTENTIAL DEVELOPMENTS 439 

low-pressure cylinder would probably have to be about three 
times that of the high-pressure cylinder, or with the same stroke 
its diameter would be 1.7 times that of the high-pressure cylin- 
der. Attempts to construct compound gas engines have been 
made in stationary types but without commercial success. The 
extra power obtainable has not justified the additional first cost 
and maintenance. It is possible, however, that more success 
may be met in the airplane engine where first cost is not of prime 
importance. The weight of the engine per horse power should 
not be increased by compounding and the weight of fuel used 
should be appreciably diminished. 

Two-cycle. — A modification of the Otto cycle which offers 
possibilities of considerable reduction in weight per horse power 
is two-cycle operation. The normal Otto cycle requires four 
strokes for the completion of the cycle of which two are used 
solely for pushing out the burned charge and taking in the new 
charge. The essential parts of the cycle are unchanged if the 
exhaust and admission processes are speeded up and made, 
in part, simultaneous by using a slightly precompressed 
charge to sweep out the exhaust gases after the exhaust pressure 
has fallen substantially to atmospheric pressure. With this 
arrangement the cycle of operation may be completed in two 
strokes, the number of cycles per minute doubled and the horse 
power almost doubled. Considerable experience with two- 
cycle engines is available from stationary practice and marine 
practice and indicates the possibility of increase of the power 
output from a cylinder of given size by from 60 to 80 per cent but 
with a falling off in efficiency of 20 per cent or more. 

There are two general methods of obtaining a precompressed 
charge, (1) by crankcase compression and (2) by the use of a 
separate compressor. With crankcase compression in a multi- 
cylinder engine, the crankcase must be divided to form a 
gas-tight compartment for each cylinder so that each piston on 
its down stroke may compress a charge which has been taken in 
during the up stroke. This arrangement is only possible in an 
engine with a single row of cylinders; it cannot be carried out in 
a Vee, W, or radial engine. It is simpler than the separate 
compressor but it limits the amount of charge which can be 
taken in to the volume sucked in during the up stroke of the 
piston and it will carry lubricating oil from the crankcase into 
the cylinder. A separate compressor should have a displacement 



440 



THE AIRPLANE ENGINE 



volume greater than that of the engine cylinder in order to send 
some scavenging air into the engine cylinder for the more com- 
plete clearing out of the exhaust gases and also to give a pressure 
at the beginning of compression which is fully up to or slightly 
above atmospheric pressure. The compression pressure required 
is about 5 lb. per square inch. One double-acting air compressor 
would be required for two engine cylinders with discharge from 
the compressor direct to the cylinders. By the use of a receiver 
a larger compressor can be made to serve a larger number of 
cylinders. A centrifugal compressor would eliminate the need 
for a receiver. 

The two-cycle engine is just beginning to be used in airplane 
service. The principal difficulties to overcome are low efficiency 
and heat trouble. If the additional weight of fuel that has to be 



Annular Exhaust 
Passa< 




Cooling Water In let 



Fig. 341. — Junkers two-cycle solid-injection engine 



carried for a long flight is equal to the saving in engine weight 
there is little advantage in the lighter weight engine except for 
short flights. The doubled number of explosions in the engine 
per unit of time increases the cylinder temperature and leads to 
serious heat difficulties with the piston. Exhaust valve troubles 
are eliminated by the use of exhaust ports uncovered by the 
piston in place of exhaust valves. There has been considerable 
experimental work in this field but with no practical results so far 
except in the case of the Junkers engine. 

The Junkers engine (Fig. 341) obtains high efficiency and 
eliminates heat trouble by departing entirely from the usual 
construction. There are six horizontal cylinders with their 
axes at right angles to the center line of the fuselage. The 
engine has two opposed pistons per cylinder and two crankshafts. 
All the pistons on each side of the engine are connected to a 



POTENTIAL DEVELOPMENTS 441 

common crankshaft. The two crankshafts are geared together 
so as to make the two pistons of any one cylinder move in or out 
simultaneously. The combustion chamber is the space enclosed 
between the two pistons when they are on their inner dead 
center. Near the outer dead centers the pistons uncover cylinder 
ports, the exhaust ports (on the left) being uncovered first and 
the air admission ports (on the right) shortly afterwards. The 
propeller shaft carries the central gear with which the gears on 
the two crankshafts mesh; a blower is operated from the pro- 
peller shaft. There are no valves and no carburetors. When 
the pistons are on their outer dead centers, air from the blower 
passes through the cylinder from right to left and clears out the 
exhaust gases. This air is compressed, while the pistons make 
their inward strokes, to a pressure of 210 lb. per square inch or 
more. Fuel is then injected into the combustion space 
through the nozzle at the bottom of the cylinder, is ignited by a 
spark plug immediately above it, and expands, driving the two 
pistons outward. The pistons are equipped with a special 
cooling device. They are made with a cavity which is partly 
filled with a heavy oil and then sealed. The oil is violently 
dashed backwards and forwards by the motion of the piston, 
absorbs heat from the piston head and carries it to the cooled 
piston sides. The efficiency of this method of cooling is shown 
by tests with thermo-elements which indicated a maximum 
temperature of the piston head of 350°F. at maximum speeds 
and loads. 

The advantages of this method of construction appear to be 
manifold. The high compression gives high efficiency, which is 
helped by the small heat loss during explosion resulting from the 
smallness of the cooling surface of the combustion chamber. The 
excellent scavenging permits higher volumetric efficiency, which 
in conjunction with higher efficiency gives a higher m.e.p. than 
is obtainable in other types. The i.h.p. is consequently more 
than twice that obtained for the same piston displacement in 
four-cycle engines and the weight is reduced to 1.5 lb. per horse 
power. The balancing of the reciprocating parts is practically 
perfect because the two pistons of each pair are at all times 
moving with equal accelerations in opposite directions; this 
condition is favorable to high engine speeds. The large size of 
the gas inlet and exhaust ports combined with positive admission 
of the air and fuel permits also high volumetric efficiency at 



442 THE AIRPLANE ENGINE 

high speeds; consequently this engine can be operated at higher 
speeds than the usual type. As the propeller is geared it can be 
run at its most favorable speed. A further feature of the engine 
is the great reduction of fire risk resulting from the direct dis- 
charge of the fuel into the engine cylinder; there is no explosive 
mixture outside the engine cylinder and less liability to fuel 
leaks. Moreover, a less volatile fuel can be burned. 

The principal apparent objection to the Junkers engine is the 
difficulty of accommodating an engine of its width in the fuselage. 
In the design shown in Fig. 341, the over-all width is 8J^ times the 
stroke, or with a 6-in. stroke the width is 4 ft. 3 in. This dimen- 
sion is exceeded in large radial engines (see p. 195). The mechan- 
ical efficiency is low on account of the blower work and of the 
gearing losses; other Junkers engines, not adapted to airplane 
use, have given mechanical efficiencies of about 73 per cent. 1 

Among the possible developments for airplane engines are 
Diesel engines, gas turbines and steam plants either turbine 
or reciprocating. The Diesel engine would be most advantage- 
ous in view of its higher efficiency and the safety and low cost of 
the fuels which it could employ. The difficulties in the way of its 
employment in airplanes in its present stage of development are 
its excessive weight and the large size of individual cylinder 
below which it has not been found practicable to go. In the 
modified form of the Hvid and similar engines, smaller size 
cylinders become practicable but the weight is still excessive. 
It is by no means certain that it will be found practicable to 
operate this cycle successfully at the high speeds necessary for 
airplane use. 

Gas turbines have been under active development for over 
fifteen years but the difficulties inherent in them have not as yet 
been overcome without sacrificing their potential efficiencies. 
Th.e principal troubles are those resulting from the high temper- 
atures to which the combustion chamber, nozzles, and buckets 
are subjected. When these temperatures are reduced, by inject- 
ing water or excess air, the efficiencies fall off. Moreover, the 
efficiencies are low unless the air is precompressed and as centrif- 
ugal compressors seldom have efficiencies above about 60 per 
cent, a large part of the power developed in the turbine is utilized 
in driving the compressor. Over-all thermal efficiencies are 
usually about 5 per cent, although an unsubstantiated value of 

1 Scott, Jour. Soc. Aut. Eng., 1917. 



POTENTIAL DEVELOPMENTS 443 

20 per cent has been claimed for a 1,000-h.p. unit. The gas 
turbine, if applied to airplane propulsion, would have to be geared 
down, probably with double reduction gear. Its simplicity and 
light weight have attracted many inventors, but there are no 
indications that it is ever likely to become practically available. 
Steam-power plants have the great advantage over internal 
combustion engines of maintaining their power at all altitudes. 
Both turbines and reciprocating engines of light weight are 
fully developed and can be regarded as immediately available 
for airplane use. The difficulties arise in connection with the 
boiler and oil burner. For efficiency the steam must be generated 
at high pressure and with high superheat, but no construction of 
boiler is known which does not entail weights which would be 
excessive for aircraft. A satisfactory kerosene (or other fuel) 
burner for operation with high rates of combustion in small 
space would also have to be developed; the fire risk from such 
apparatus would probably be considerable. And finally, the 
efficiency of the best steam plants is not nearly so high as that of 
existing airplane engines. It seems very unlikely that steam 
plants will ever be employed in aircraft. 



INDEX 



A. B. C. engines, (table), 194 
Acceleration, in carburetors, 271 

force, (see Inertia Force) 
Acetylene, as fuel, properties of, 225 
Aerofoil, (see Wings) 
Air compression, power absorbed in, 
397 
temperature rise in, 395, 400 
cooling, 344 
cycle efficiency, (def. and table), 

17 
densities, at altitudes, (curves), 

366 
flow, pulsating, 254 
flow of, orifice coefficients for, 243 
through venturi tube, 246- 
250 
fuel ratio, determination of 
strength of, 242 
influence of charge dilution 
on optimum value of, 262 
influence of, on engine per- 
formance, 260 
optimum values of, 261 
variation with air density of, 
259 
intakes, 416 
pumps, 292 

starting, (see also Starting), 429 
temperature, at altitude, (curves), 
365 
influence of, on capacity, 35 
on engine power, 33, 35 
on thermal efficiency, 35 
weight flow of, (chart), 249 
Alcogas, 242 

Alcohol, as fuel, properties of, 224 
effect on detonating pressure, 239 
mixtures with gasoline, 242 
Alloys (see Steels, and Aluminum 
Alloys) 



Altitude, air density at, (curves), 366 
control of carburetor, 268 
effect of, on radiator, 365 
influence of, on engine power, 

386-389 
temperatures at, (curves), 365 
Aluminum alloys, 118 

for pistons, 135 
American engines, description of, 

80-98 
Austro -Daimler engine, cylinder of, 
131 
dimensions of, 73, 124 
inertia forces and bearing loads, 

59 
water pumps of, 372 
weights of parts, 78 



B. R. 2 engine, description of. 190 
Balancing, devices for, 58 
in radial engines, 206 
in rotary engines, 206 
of reciprocating parts, 55 
of rotating parts, 54 
Ball and ball carburetor, description 

of, 283 
Battery ignition systems, (see Igni- 
tion Systems) 
Battery, of Liberty engine, 316 
Basse -Selve carburetor, description 
of, 284 
engine, compression release of, 428 
cylinder dimensions of, 124 
dimensions of, 73 
inertia forces and bearing loads 

of, 59 
lubricating system of, 341 
oil pumps of, 341 
valve gear of, 171 
valves of, 157 
weights of parts, 78 



445 



446 



INDEX 



Baume scale, conversion table for, 

219 
Bayerische Motoren Werke carbu- 
retor, description of, 285 
Bearings, 144 
ball and roller, in radial engines, 

207 
crankshaft, dimensions of, 74 
friction work of, 328 
loads on, 59, 76, 327 
oil grooving of, 337 
Benz engine, compression release of, 
427 
connecting rods of, 130, 141 
cylinder of, 124, 129 
description of, 110 
dimensions of, 73 
fuel pump of. 294 
intake manifold of, 420 
lubrication system of, 335 
oil pump of, 336 
performance curves of, 111 
piston of, 136 
propeller hub of, 148 
valve gear of, 172 
weights of parts, 78 
Benzene, (see Benzol) 
Benzol, as fuel, properties of, 223 
Bijur starting system, 432 
Bosch magneto, 302 
Brake mean effective pressure, (def.), 

25 
Bugatti engine, description of, 93-98 
performance curves of, 98 
spark adjustment in, 306 
water pumps of, 372 



Cams, 163 

followers for, 164 
Camshafts, dimensions of, 72 
Capacity of engine, 26, 29 
influence of fuels on, 240 
variation with air temperature, 
33, 35 
compression ratio, 37, 392 
engine speed, 37, 392 
jacket-water temperature, 37 
mixture strength, 33 



Carbon dioxide, dissociation of, 20 
Carburetors, 245-289 
acceleration in, 271 
air discharge coefficients of, 252 
altimetric compensation of, 267 

control of, 268 
atomization in, 271 
construction of, 272 
dimensions of, 74 
float arrangements in, 272 
flooding of, 272 
idling device in, 271 
intakes for, 272 
mixture characteristics of, 258 
performance of, 264 
pressure drop in, 253 
strainers for, 290 
viscous flow type, 267 
weights of, 78 
Castor oil, 333 
Central power plants, 384 
Choke, (see also Venturi Tube), 250 
Christensen system of starting, 431 
Claudel carburetor, description of, 

276 
Clerget engine, description of, 189 
effect of compression ratio on 

horsepower of, 393 
performance curves of, 190 
Combustion, air required for, 228 
effect of turbulence on rate of, 235 
higher and lower heats of, (def.), 

214 
products of, (table), 216 
velocity of propagation of, 232 
Compound engines, 438 
Compression, ratio of, (def.), 14 
variable, 438 
pressures, effect of alcohol on, 239 
effect of toluene on, 239 
maximum allowable, 236. 237 •$ 
variation with compression 
ratio, 16 
engine speed, 16 
ratio, 66, 72 

influence of, on capacity, 37 
on efficiency, 434 

engine power at high 
altitudes, 392 



INDEX 



447 



Compression, release, 427 
Compressors, (see Air Compression) 
Connecting rod, rods, 138 

assembly, articulated, 203 

dimensions of. 74 

for rotary and radial engines, 

203 
materials for, 116 
slipper assembly, 204 
stresses in articulated, 143 
weights of, 76, 78 
Cooling fins, 345 
systems, 344 
anti-freeze solutions for, 375 
piping for, 375 

pumps for, (see Water Pumps) 
typical examples of, 377 
water for, 375 
Cosmos engines, (table), 195 
induction chamber of, 196 
"Jupiter," balancing of, 207 
performance curves of, 196 
Counterweights, 146 
Crankcases, 148 
cooling of, 150 
weights of, (table), 78 
Crank pins, loads on, 76 
Crankshafts, 143-146 
balancing of, 143 
dimensions of, 74 
materials for, 115 
of radial engines, 211 
strength of, 146 
weight of, 78 
Curtiss engines, crankshaft balanc- 
ing of, 144 
cylinders of, 124, 127 
description of, 86-90 
lubricating systems of, 334 
performance curves of, 90, 92 
type K, 86 
Cylinders, air-cooled, 200, 345 
materials for, 349 
temperature of, 347 
arrangements, size and propor- 
tions of, 61 
attachment to crank case, 125 
liners, material of, 115 
lubrication of, (see Lubrication) 



Cylinders, offset of, 48 
thickness of, 123 
types of construction, 122 
weights of, 78 



Detonation, 236 
Diesel engines, 442 
Dilution of charge, influence of 
compression pressure on, 
263 
influence on engine perform- 
ance, 393, 434 
Dimensions, engine, (table), 65 
of American and German engines, 

(table), 72 
overall, of engines, (table), 79 
Dissociation, of carbon dioxide, 20 

of water vapor, 20 
Distillation curves, 241 
Distributor, Bosch, 304 

Dixie, 305 
Dixie magneto, 302 
Double-rotary engines, 176, 191 
Drag, of wing, 2 
Duralumin, 121 



E 



Engine speed, influence of, on 

capacity, 37 
English engines, descriptions of, 

98-107 
Ether, as fuel, properties of, 226 
Exhaust gas, composition of, 244 
turbines, (see also Supercharg- 
ing), 398, 438 
manifolds, (see Manifolds) 
mufflers, (see Mufflers) 
Explosion phenomena, (see also 
Fuels, Combustion), 231 
intervals, in multicylinder engines, 

63 
limits, effect of carbon dioxide 
dilution on, 262 
Explosive mixtures, 212 

properties of, (table), 216 
wave, 236 



448 



INDEX 



Fiat engine, description of, 107 
lubrication system of, 335 
valves of, 174 
Firing order, 75, 325 
Flash point, of oils, 330 
Flight, power available for, 8 

power required for, 1 
Friction, laws of, 327 
loss, at piston, 24 
in engine, 24 
Fuel, fuels, 212-226 

air ratio, {see Air-fuel Ratio) 

air required for combustion of, 

(table), 216 
detonating compression pressures 

of, (table), 237 
explosive limits of air-fuel mix- 
tures, (table), 232 
flow through jets, 254-258 
heats of combustion of, (table), 

217 
ignition temperatures of, 233 
influence on capacity of, 240 
influence of temperature on 

fluidity, (curves), 257 
hydrocarbon, classification of, 213 
minimum vaporization tempera- 
ture of, 231 
properties of, 212, 215 
pumps, 289, 292 
weights of, 78 
specific heats of, 229 
specific volumes of saturated 

vapors of, 229 
systems, 289-294 

for supercharging engine, 412 
tanks, 290 

temperature drop due to vapori- 
zation of, 230 
vapor pressures of, 229 
viscosity of, 257 



Gas turbines, {see also Exhaust Gas 
Turbines), 442 
velocity, through valves, 152 



Gasoline, {see also Fuel), 217-223 
blended casing-head, 218 
calorific value of, 223, 241 
cracked, (synthetic), 218 
distillation curves of, 221 
mixtures with alcohol, 242 
specifications for, 219 
"straight" refinery, 218 
tests for, 222 
volatility of, 220 

Geared propeller drives, 378-384 

Gears, heating of, 384 
pressure between teeth of, 383 
stresses in, 383 

Gnome engine, description of, 185 
torque of, 53 

Gudgeon pin, {see Piston Pin) 



H 



Hall-Scott engine, cylinder dimen- 
sions of, 124 
description of, 90 
exhaust manifold of, 420 
lubrication system of, 335 
performance curves of, 95 
Heat balance, of airplane engine, 
38 
dissipation, from air-cooled cylin- 
ders, 345 
transfer, in radiator core, 352 
Hispano -Suiza engine, air pump of, 
292 
cylinder of, 126 
description of, 82 
dimensions of, 73 
effect of compression ratio on 

horsepower of, 393 
hand starting mechanism of, 430 
influence of air density on 

performance, 388 
lubrication system of, 334 
performance curves of, 86 
propeller gears of, 378 
propeller hub of, 147 
starting torque of, 427 
test results of, 31 
valves of, 173 
weights of parts of, 78 



INDEX 



449 



Horse power (see also Capacity) 

required for flight, 4 
Hydrogen, as fuel, properties of, 225 



Idling device, in carburetors, 271 
Ignition, 295-326 

assemblies, weights of, 78 
spark advance, 306 
systems, battery, 297, 312 
comparison of, 324 
cycle of operations in, 307 
self-sustaining battery, 313 
temperature of, 233 
Indicated thermal efficiency, maxi- 
mum obtainable, 237 
Indicator cards, actual, 15 

negative pumping loop, 23 
theoretical, 13 
Induction coil, 296 
Inertia factors, (table), 42 
. forces, in Austro-Daimler, Basse- 
Selve and Liberty engines, 
59 
in radial engines, 52 J 
primary, 55 
secondary, 55 
of reciprocating parts, 41 
Intake manifolds, (see Manifolds) 



Jacket-water temperature, influence 

of, on capacity, 37 
Jets, discharge coefficients of, 255 

flow through, 254 
Junkers two-cycle fuel-injection 
engine, 440 



K 



Kerosene, as fuel in airplane engines, 
436 



Lanchester balancer, 58 

LeRhone carburetor, description of, 

288 

29 



LeRhone engine, description of, 185 
effect of compression ratio on 

horse power, 393 
oil pumps of, 342 
performance curves of, 187 
Liberty engine, air intake for, 417 
battery ignition system of, 

313 
breaker mechanism of, 318 
centrifugal oil cleaner of, 333 
connecting rods of, 141 
cylinder of, 124, 130 
description of, 80 
dimensions of, (table), 72 
distributors of, 318 
electric starting system for, 

432 
generator, 315 
heat balance, 39 
indicator card, 40 
inertia forces and bearing loads, 

59 
influence of air density on 

performance of, 388 
intake manifolds of, 419 
lubrication system of, 334 
method of starting, 425 
oil pumps of, 339 
performance curves of, 31, 83, 

239 
piston of, 138 
starting torque of, 427 
torque of, 43 

valve action of, (curves), 168 
valve of, 172 
water pumps of, 371 
weights of parts, (table), 78 
Lift, on a wing, 1 

Lubricating oils, properties of, 
(table), 331 
reclaiming of, 332 
specifications for, 330 
tests of, 330 
viscosity of, 329 
Lubrication, 327-343 
methods of, 333 
of cylinders, 328 
of radial engines, 211 
oil consumption for, 338 



450 



INDEX 



M 



Magneto, 297-307 
armature flux, 308 
typical constants for, 307 
for unequal firing intervals, 300 
inductor type, 300 
secondary voltage in, 309 
speed of, 306 
starting, 425 
Manifolds, exhaust, 421 
ejector effect in, 422 
intake, 417-421 

influence of, on engine perform- 
ance, 436 
preheating of, 419 
pressure drop in, 28 
temperature change in, 418 
Master carburetor, description of, 

283 
Materials, for special engine parts 
(table), 118 
properties of, (table), 114 
Maybach carburetor, description of, 
286 
engine, cylinder of, 129 
dimensions of, 124 
description of, 113 
dimensions of, (table), 73 
fuel pump of, 293 
lubrication system of, 336 
performance curves of, 113 
pistons of, 136 
starting mechanism of, 426 
valve action of, (curves), 167 
valve gear of, 166 
water pumps of, 374 
weights of parts, (table) 78 
Mechanical efficiency, 23 
Mean effective pressure, (table) 67 
brake, 25 

maximum obtainable, 237, 
238 
Mercedes engine, air pumps of, 
292 
dimensions of (table), 73 
weights of parts (table), 78 
Miller carburetor, description of, 
282 



Mixture strength (see also Air-fuel 
Ratio) 
influence of, on engine perform- 
ance, 22, 32 

Moss turbo-supercharger, 411 

Mufflers, exhaust gas, 422 



N 



Napier "Lion, " connecting rods of, 
142 
description of, 103 
lubrication system of, 335 
performance curves of, 106 



O 



Odier portable starter, 433 
Offset cylinders, 48 
Oil pumps, 336, 338 

performance curves of, 340 

weights of (table), 78 
Oil sumps, 150 
Oil tanks, 343 
Oversized engines, 390-394 
Orifice, discharge coefficients for 

sharp-edged, 243 
Otto cycle, 13 

efficiency with variable specific 
heat, 19, 21 
Oxygen, use in engine at altitudes, 
413 



Packard engine, cylinder head of, 
131 
description of, 80 
dimensions of (table), 72 
performance curves of, 84 
weights of parts of (table), 78 

Parasite resistance, 3 

Performance curves, Benz engine, 
111 
Bugatti engine, 98 
Clerget engine, 190 
Cosmos "Jupiter" engine, 196 
Curtiss engine, 90, 92 
Hall-Scott engine, 98 



INDEX 



451 



Performance curves, Hispano-Suiza 
engine, 86 
LeRhone engine, 187 
Liberty engine, 83 
Maybach engine, 113 
Napier "Lion" engine, 106 
Packard engine, 84 
Salmson engine, 199 
Siddeley "Puma" engine, 106 
Rolls-Royce engine, 102 
Pipes, for fuel system, 290 
Piston, Pistons, dimensions of, 
(table), 72 
displacement, per horsepower. 

(table), 67 
divided-skirt, 135 
friction of, 133 
material of, 132 
pin, 138 

dimensions of, (table), 74 
loads on, (table), 76 
rings, 137 

dimension of, 74 
slap of, 134 
slipper, 133 
speed, (table), 66 
weights of, (table), 76, 78, 166 
working temperatures of, 132 
Pitch, of propeller, (see Propeller 

Pitch of) 
Power, (see Horsepower and Capacity) 
available, for flight, 8 
required, for flight, 1 
Priming, of engine, 423 
Pumps, fuel, 289, 292 
Pressure, Pressures, in ideal (air) 
cycle, 14 
drop, in intake manifold, 28 

past valves, 152 
mean effective, 25 
on bearings, 144 
on crankpin, 43 
on piston, 40 
Propeller, 5 
coefficients, 7 
efficiency of, 6 
geared, (see Geared Propeller 

Drives) 
hubs, 147 



Propeller hubs, weights of, (table), 78 
pitch, 6 
pitch ratio, 6 
slip, 6 

speed, (table), 66 
thrust, 6 

thrust horse power of, 6 
torque, 6 
torque horse power of, 6 

R 

Radial engines, 176, 193 

air-cooled, dimensions of, 

(table), 66 
ball and roller crankpin bearings 

in, 207 
crankshaft of, 211 
details of, 200 
firing order of, 178 
lubrication of, 211 
number of cylinders of, 178 
overall dimensions of, (table), 79 
unbalanced forces in, 56 
valve operation in, 211 
water-cooled, dimensions of, 
(table), 68 
Radiators, 351-370 

constants of, (table), 362 
construction of complete, 368 
cores of, 351 

dimensions of, 359 
effect of position on performance 

of, 363 
figure of merit of, 355, (table), 360 
head resistance of, 354, 356, 

(table), 360 
heat transfer in, 352, 354, (table), 

360 
horse power absorbed by, 354, 

(table), 360 
limiting temperatures in, 363 
masking of, 366, 368 
mass flow of air in, (def.), 352 
obstructed, 352 

on lighter-than-air machines, 367 
performance of, 358, 360 

at altitudes, 365 
resistance to water flow in, 364 
selection of, 356 



452 



INDEX 



Radiators, size of, 361 
water flow through, 363 
yawing of, 368 
Rateau supercharger, 410 
Reduction gearing, (see Geared Pro- 
peller Drives) 
Renault engine, connecting rods of, 
142 
cylinder dimensions of, 124 
propeller gears for, 378 
Resistance, parasite, 3 
of plane, 4 
of wing, 2 
Revolutions, of typical engines, 

(table), 66 
Ricardo system, of supercharging, 

401 
Rocker arms, 171 

Rolls-Royce engine, description of, 98 
dimensions of, (table), 101 
performance curves of, 102 
propeller gears for, 380, 382 
Rotary engines, 176, 179-193 
balance of, 57 
details of, 200 
dimensions of, (table), 66 
firing order of, 178 
number of cylinders of, 178 
overall dimensions of (table), 79 
torque of, 50 

S 

Salmson engine, description of, 198 

performance curves of, 199 
Saturation temperature of air-fuel 

mixtures, (table), 231 
Side thrust, 48 
Siddeley "Puma" engine, 11 
connecting rod of, 140 
description of, 105 
performance curves of, 106 
valves of, 173 
Siemens-Halske double-rotary en- 
gine, 192 
Slip, of propeller, 6 
Spark advance, (see Ignition) 
gaps, 310 
plugs, 319-324 

causes of failure of, 320 



Spark plugs, construction of, 
322 
cracking of insulators of, 321 
dimensions of, 76 
location and number of, 324 
Sparking voltages, 311 
Specific heats, at constant volume, 

(table), 20 
Springs, safe loads and deflections 

of, 170 
Starting, 424-433 

air compressors for, 430 
by hand, 424 
Christensen system, 431 
compressed air for, 429 
compression release for, 427 
electric systems for, 432 
magneto, 425 

mechanisms, dog clutch for, 428 
integral, 429 
safety devices for, 429 
Motor-Compressor Company sys- 
tem, 430 
portable crankers for, 433 
torque, 426 

use of hydrogen for, 425 
Steam-power plants, 443 
Steels, 116-117 

for exhaust valves, 160 
Stewart-Warner carburetor, air dis- 
charge coefficients of, 252 
mixture^characteristics of, 258 
performance curves of, 265 
Storage cells, (see also Battery), 312 
Strainers, for fuel systems, 290 
Stroke-bore ratios, 66, 72 
Stromberg carburetor, description 
of, 278 
mixture characteristics of, 

(curves), 258 
performance curves of, 266 
Sturtevant engine, cylinder of, 128 

supercharger for, 404 
Supercharged engines, 394-413 
explosion relief valve for, 408 
fuel supply system for, 412 
power developed by, 394 
relief valve of, 413 
throttle valve for, 408 



INDEX 



453 



Superchargers, (see also Air Com- 
pression), 397 

Brown-Boveri, 407 

coupling for, 407 

exhaust gas turbine, 398, 408 

Moss turbo-, 411 

Rateau, 410 

Schwade, 406 

Sturtevant, 404 
Supercharging, 368-415 

centrifugal compressors for, 403 

efficiency of, compressor, 400 
exhaust gas turbine, 400 

gearing of compressors for, 403 

influence on plane performance, 
415 

methods, 401 

multi-stage compressors for, 404 

reciprocating compressors for, 403 

Ricardo system of, 401 

Roots positive blower for, 403 



Torque, in rotary engines, 50 
of engine, 26, 67 

at starting, 426 
of propeller, 6 

ratio of maximum to mean, 46 
variation with number of cylin- 
ders (table), 47 
Torsional vibration, of crank-shaft, 

146 
Turbines, exhaust gas, (see Exhaust 
Gas Turbines, Super- 
charging) 
Turbo-superchargers, (see Super- 
chargers) 
Turbulence, 235 
Turning moment, (see also Torque), 

40 
Two-cycle engines, 439 

U 

Unbalanced forces, magnitude of, 56 
neutralizing of, 58 
periodic, effects of, 57 



Tangential factors, (table), 44 
Tanks, fuel, 290 ' 

oil, 343 
Tappet clearance, adjustment of, 172 
Temperatures, in ideal (air) cycle, 
14 
of jacket-water, influence on en- 
gine capacity, 37J 
Test results, correction to standard 

conditions, 32 
Thermal efficiency, 25 

effect of mixture strength on, 22 
maximum observed, 21 
variation with air temperature, 
35 
mixture strength, 33 
throttling, 36 
Throttling, influence of, on thermal 

efficiency, 36 
Thrust, of propeller, 1, 6 
Timing (see Valve Timing) 
Toluene, effect on detonating tem- 
perature, (table), 239 
value (def.), 238 
Torque, at engine crankshaft, 43 



Valve, Valves, automatic throttle, 391 
causes of failure of, 160 
effect of lift and diameter on 

engine capacity, 158 
exhaust, 156 

dimensions of, 66, 72 

temperature of, in air-cooled 
cylinder, 349 
gears, 162 

weights of, (table), 78 
inlet, dimensions of, 66, 72 

number per cylinder, 66, 72 
lift of, 151 

materials of, 159, 161 
operation, in radial engines, 211 
ports, 125 

pressure drop past, 152 
seats, 125 
springs, 168 

retainers for, 170 

tension of, (table), 72 
stem guides, 125 
temperatures of, 159 
timing of, 172, 173, 175 



454 INDEX 

Valves, weights of, (table), 76 Water cooling, 350 

Variable-compression engine, 438 pumps, 370 

stroke engine, 437 horse power required for, 370 

Vee engines, angle of, 64 performance curves of, 373, 374 

connecting rods of, 139 weights of, (table), 78 

dimensions, (table), 70 vapor, dissociation of, 20 

overall dimensions of, (table), 79 Weights, of engines, 60, 67, 78 

unbalanced forces in, 56 of engine parts, (table), 78 

Venturi tube, 251 of reciprocating parts, (table), 

discharge coefficients of, 252 76 

Vertical engines, dimensions of, of rotating parts, (table), 76 

(table), 68 of water in engine, (table), 78 

overall dimensions of (table), 79 Wings, characteristics of, 2 

unbalanced forces in, 56 Wiring systems, 325 

Vibration, {see Torsional Vibration) Wright engine, description of, 84 

Viscosimeter, 329 Wrist pin, {see Piston 'pin) 
Viscosity, measurement of, 329 

of fuels, 257 Z 
of oils, 329 

Volatility, {see also Distillation), 220 Zeitlen engine, description of, 437 

Volumetric efficiency, 27, 28 Zenith carburetor, air discharge 

coefficients of, 252 

W description of, 272 

W engines, arrangements of, 65 mixture characteristics of, 258 

dimensions of, (table), 70 performance curves of, 264 



